FLIGHT MANUAL
COMMANDERS ARE RESPONSIBLE
FOR BRINGING THIS TECHNICAL
PUBLICATION TO THE ATTENTION
Of ALL AIR FORCE PERSONNEL
CLEARED FOR OPERATION OF
AFFECTED AIRCRAFT.
PUBLISHED UNDER AUTHORITY OF THE
SECRETARY OF THE AIR FORCE.
THIS MANUAL IS NOT COMPLETE WITHOUT
CONFIDENTIAL SUPPLEMENT T.O 1M9H-1A.
THIS CHANGE INCLUDES SAFETY OF FLIGHT SUPPLEMENTS
THROUGH -1Y. SEE BASIC INDEX, T.O. 0-1-i, AND WEEKLY INDEX, T.O
FOR CURRENT STATUS OF SAFETY OF FLIGHT SUPPLEMENTS.
LATEST CHANGED PAGES SUPERSE
THE SAME PAGES OF PREVIOUS D/I
Insert changed pages into basic
publication. Destroy superseded pages.
31 OCTOBER 1058
CHANGED 13 FEBRUARY 1959
AIR FORCE, Kerr Litho, Culver City, Cslif., 4/12/59-1450 (Northrop)
TO. 1F-89H-1
Reproduction for nonmilitary use of the information or illustrations contained in this publication is not per¬
mitted without specific approval of the issuing service (BuAer or USAF). The policy for use of Classified
Publications is established for the Air Force in AFR 205-1 and for the Navy in Navy Regulations, Article 1509.
LIST OF EFFECTIVE PAGES
INSERT LATEST CHANGED PAGES. DESTROY SUPERSEDED PAGES.
NOTE: The porrion of the text affccrcd by the changes is indicated
by a vertical line in the outer margins of the page.
TOTAL NUMBER OF PAGES IN THIS PUBLICATION IS 376, CONSISTING OF THE FOLLOWING:
Page Date of Latest
No. Issue
*Title Page.13 February 1959
*A . 13 February 1959
i through iv ..Original
1-1 through 1-3. .Original
*1-4.13 February 1959
1-5 *. -.. *...Original
* 1-6.13 February 1959
1- 7 through 1-66., *... .Original
2- 1 .. .Original
*2-2 through 2-5.13 February 1959
2- 6 through 2-44. Original
3- 1 through 3-2. .. . Original
*3-3 . 13 February 1959
3-4 through 3-11.Original
*3-12 . 13 February 1959
3-13 through 3-18. .Original
*3-18A through 3-19.. 13 February 1959
3-20 through 3-28. ..........Original
*3-29 . .13 February 1959
3- 30 through 3-38.Original
4- 1 through 4-3............ .Original
*4-4 ................ 13 February 1959
4- 5 through 4-32.. .Original
5- 1 through 5-16....Original
6- 1 through 6-8... Original
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6- 10 through 6-18. ......... .Original
7- 1 through 7-6.Original
*8-1 through 8-2..13 February 1959
8- 3 through 8-6..Original
9- 1 through 9-22........... .Original
A-i through A-1IG. ....... * .Original
1 . Original
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3 through 5. Original
*6 ... 13 February 1959
7 ..Original
*8 . 13 February 1959
9 through 10.. ...... Original
* The asterisk indicates pages changed, added, or deleted, by the current change.
ADDITIONAL COPIES OF THIS PUBLICATION MAY BE OBTAINED AS FOLLOWS: D-l
USAF
USAF ACTIVITIES.—In accordance with Technical Order 00-5-2.
NAVY ACTIVITIES.—Submit request to nearesr supply point listed below, using form NavAer'140: NASD. Philadelphia. Pa.;
NAS, Alameda, Calif.; NAS, Jacksonville, Fla.; NAS, Norfolk. Va + ; NAS, San Diego, Calif,; NAS, Seattle, Wash.; ASD,
NSC, Guam.
For listing of available material and details of distribution see Naval Aeronautics Publications Index NavAer 00-500.
Changed 13 February 1959
T.O. 1F-89H-1
*
iMumnr
■' • ■> '-VV’U
mu or coHTtms
Description* •.. *1-1
Normal Procedures a******************** 2-1
III Emergency Procedures ******••***••*•••* 3-1
IV Auxiliary Equipment **•••***•*»«•••*•** * 4-1 *
m
If
V Operating Limitations••••••••■••••■•«••• 5-1
KSSfc
Ifi® -
VI Flight Characteristics *•••••••**••»»•••*• 6-1
y\\ Systems Operation•**#••***••*•••*••*** 7-1
^S^HRJUEilUEXS^fim
will Crew Duties a************************ 8-1
Mffe
IX All-Weather Operation ••*•***#«*••***#+« 9-1
I Performance Data * *••••**»••*•••**••«* A-l
Alphabetical *••*••*«*•****•••***•**** *X -1
Prefer to t.o. if-ssh-ia for additional information.
H-2D
i
TO. 1F-89H-I
BECAUSE OF THE RIGID REQUIREMENTS IMPOSED ON INTERCEPTOR CREWS OPERATING
THE HIGH PERFORMANCE FIGHTERS OF THE JET AGE, LADY LUCK CANNOT BE RELIED
UPON FOR COMPLETION OF A SUCCESSFUL MISSION. THEREFORE, IT IS MANDATORY
THAT COMPLETE UNDERSTANDING OF FLIGHT CHARACTERISTICS AND OPERATING TECH¬
NIQUES FOR THE HIGHLY COMPLEX SYSTEMS BE MAINTAINED AT THE HIGHEST LEVEL.
SCOPE
This manual contains all the information necessary
for safe and efficient operation of the F-89H. These in¬
structions do not teach basic flight principles, but are
designed to provide you with a general knowledge of the
airplane, its flight characteristics, and specific normal
and emergency operating procedures. Your flying expe¬
rience is recognized, and elementary instructions have
been avoided.
SOUND JUDGMENT
The instructions in this manual are designed to pro¬
vide for the needs of a crew inexperienced in the opera¬
tion of this airplane. This manual provides the best
possible operating instructions under most circum¬
stances, but it is a poor substitute for sound judgment.
Multiple emergencies, adverse weather, terrain, etc, may
require modification of the procedures contained herein.
PERMISSIBLE OPERATIONS
The Flight Manual takes a "'positive approach” and
normally tells you only what you can do. Any unusual
operation or configuration (such as asymmetrical load¬
ing) is prohibited unless specifically covered in the Flight
Manual. Clearance must be obtained from ARDC before
any questionable operation is attempted which is not
specifically covered in the Flight Manual.
STANDARDIZATION
Once you have learned to use one Flight Manual,you
will know r how to use them all—closely guarded stand¬
ardization assures that the scope and arrangement of
all Flight Manuals are identical.
ARRANGEMENT
The manual has been divided into ten fairly inde¬
pendent sections, each with its own table of contents. The
objective of this subdivision is to make it easy both to read
the manual straight through when it is first received and
thereafter to use it as a reference manual. The indepen¬
dence of these sections also makes it possible for the user
to rearrange the manual to satisfy his personal taste and
requirements. The first three sections cover the minimum
information required to get the airplane safely into the air
and back dow r n again. Before flying any new airplane
these three sections must be read thoroughly and fully un¬
derstood. Section IV covers all equipment not essential to
flight but which permits the airplane to perform special
functions. Sections "V and VI are self-explanatory. Section
VII covers any technique or theory of operation which
may be applicable to the particular airplane in question.
The experienced pilot will probably be aw are of most of
the information in this section but be should check it for
any possible new- information. The contents of the remain¬
ing sections are fairly evident.
YOUR RESPONSIBILITY
These Flight Manuals are constantly kept current
through an extremely active revision program. Frequent
conferences with operating personnel and constant re¬
view of UR’s, accident reports, flight test reports, etc, as¬
sure inclusion of the latest data in these manuals. In this
regard, it is essential that you do your parti If you find
anything you don't like about the manual, let us know r
right away. We cannot correct an error if its existence is
unknown to us.
II
T*0* 1F-89H-1
PERSONAL COPIES, TABS, AND BINDERS
In accordance with the provisions of AFR 5-13,
flight crewmembers are entitled to have personal copies
of the Flight Manual* Flexible loose leaf tabs and binders
have been provided to hold your personal copy of the
Flight Manual* These handsome simulated leather hind¬
ers will make it much easier for you to revise your man¬
ual as well as to keep it in good shape* Tabs and binders
are: secured through your local materiel staff and con¬
tracting officers*
HOW TO GET COPIES
If you want to be sure of getting your manuals on
time, order them before you need them. Early ordering
will assure that enough copies are printed to cover your
requirements* Technical Order 0-5-2 explains how to
order Flight Manuals so that you automatically will get
all revisions, reissues, and Safety of Flight Supplements*
Basically, all you have to do is order the required quan¬
tities in the Publication Requirements Table (T.O. 0-3-1).
Talk to your Senior Materiel Staff Officer—it is his job to
fulfill your Technical Order requests. Make sure to estab¬
lish some system that will rapidly get the books and Safe¬
ty of Flight Supplements to the flight crews once they
are received on the base*
SAFETY OF FLIGHT SUPPLEMENTS
Safety of Flight Supplements are used to get infor¬
mation to you in a hurry. Safety of Flight Supplements
use the same number as your Flight Manual, except for
the addition of a suffix letter. Supplements covering loss
of life will get to you in 48 bourse those concerning ser¬
ious damage to equipment will make it in 10 da'ys. You
can determine the status of Safety of Flight Supplements
by referring to the Index of Technical Publications
(T.O* 0-1*1) and the Weekly Supplemental Index (T.G*
0-1-1 A)* This is the only way you can determine whether
a supplement has been rescinded* The title page of the
Flight Manual and title block of each Safety of Flight
Supplement should also be checked to determine the ef¬
fect these publications may have on existing Safety of
Flight Supplements. It is critically important that you re¬
main constantly aware of the status of all supplements*
You must comply with all existing supplements but there
is no point in restricting the operation of your airplane
by complying with a supplement that has been replaced
or rescinded.
If you have ordered your Flight Manual on the Pub¬
lications Requirements Table, you automatically will re¬
ceive all supplements pertaining to your airplane.
Technical Order 0-5-1 covers some additional informa¬
tion regarding these supplements*
WARNINGS, CAUTIONS, AND NOTES
For your information, the following definitions ap¬
ply to the "Warnings/* "Cautions," and "Notes” found
Operating procedures, prac¬
tices, etc , which will result
in personal injury or loss of
life if not carefully followed *
Operating procedures, prac -
tices, etc , which if not strict¬
ly observedi will result in
damage to equipment .
Operating procedures r condi*
tions f etc r which it is essen¬
tial to emphasize *
throughout the manual :
r\
CAUTION
Note
Airplanes having different or additional
systems and equipment have been
group coded to avoid listing of airplane
serial numbers * The groups* with
the airplanes they include* are
as shown below, right;
II.S. A lit FORCE MI9H-S^<L
A.F. SERIAL NO. AF54-416
Group 1
AF5126I THROUGH AF54-320
X
Group 5
AF54-321 THROUGH AF5U4J6
H'3C
111
COMMENTS AND QUESTIONS REGARDING ANY PHASE OF THE FLIGHT MANUAL
PROGRAM ARE INVITED AND SHOULD BE ADDRESSED TO COMMANDER, OGDEN
AIR MATERIEL AREA, HILL AIR FORCE BASE, UTAH, ATTENTION: WCLOD-3ID.
AIRFORCE 7
TO. 1F-89H-1
Section I
KSCmrtOH
T&ELi CONTENTS
The Airplane.*.* * * , , 1-1
Engines.... , , , ,.. 1-2
Afterburner System... * . . 1-9
Oil Supply System * ..,.*. 1-17
Fuel Supply System ....... . 1-17
Electrical Power Supply Systems. 1-23
Hydraulic Power Supply System. 1-33
Flight Control System. 1-35
Sideslip Stability Augmenter System ......... 1-40
Wing Flap System ... 1-41
Speed Brake System .. 1-41
Landing Gear System . .. 1-42
Nose Wheel Steering System .. 1-47
Brake System... 1-48
Instruments .. 1-48
Emergency Equipment .. 1-52
Canopy .. 1-54
Ejection Seats .. T-56
Auxiliary Equipment. 1-66
THE AIRPLANE.
The Northrop F-89H airplane is a two-place, mid¬
wing, jet-propelled, all-weather fighter interceptor
designed to operate at high speeds and high altitudes.
The airplane's function is to locate, intercept, and
destroy enemy aircraft by day or night, under ail con¬
ditions of weather. The crew consists of a pilot and a
radar observer. For maximum efficiency, the radar
equipment is operated by the observer, thus allowing
the pilot to devote his full attention to flying. This
division of duties results in higher combat effective¬
ness. The pilot and radar observer have individual
cockpits with ejection seats and automatic heating
and pressurizing facilities. The tandem cockpits are
enclosed by a single jettisonable canopy. The airplane
is powered by two turbojet engines with afterburners.
The flight control surfaces are fully powered by
two independent hydraulic systems. "Teel,” which
would otherwise be absent in a powered control sys¬
tem, is supplied artificially to the control stick and to
the rudder pedals by springs. Additional elevator
"feel” is supplied by a control force bellows system
and a "G” operated bobweight. Another unusual fea¬
ture not found on other combat airplanes is the com¬
bination of ailerons and speed brakes. Each aileron is
composed of a leading edge section and two movable
aft surfaces, one above the other, hinged at their for¬
ward edges. These two surfaces can be opened to any
desired angle, up to an included angle of 120 degrees,
to function as a speed brake. The left and right speed
brakes operate simultaneously. Pylons under the wings
carry jettisonable fuel tanks.
AIRPLANE DIMENSIONS.
Refer to figure 1-2 for dimensions of this airplane.
AIRPLANE GROSS WEIGHTS.
The design gross weight is approximately 39,500
pounds and the maximum gross weight is approxi¬
mately 47,400 pounds. See figure 5-6 for exact gross
weights.
ARMAMENT.
Standard armament consists of 2.75 folding fin aerial
rockets and GAR-1 missiles. For detailed information
on armament, refer to T.O, 1F-89H-1A, a confidential
supplement to this publication.
1-1
Section I
T.O. TF-89H-1
Figure 1-2.
ENGINES.
The airplane is powered by two J35-35 axial-flow
turbojet engines equipped with afterburners and re¬
tractable air inlet screens. Some airplanes have J3 5 -47
inner combustion liners installed on J35-35 engines.
Engines so modified have been reidentified as J35-35A
by restamping the engine nameplate and making an ap¬
propriate notation in DD Form 781. On the front
of each engine are mounted all accessories driven by
the engine shaft, including engine fuel pump, oil
pump, engine fuel control, hydraulic pump, starter gen¬
erator, 28*volt d-c generator or alternator, and tachom¬
eter generator. Air enters through the engine airscoop
and is progressively compressed through 11 stages in the
axial-flow compressor. (See figure 1-5.) A portion of
the eleventh stage compressor air is used to pressurize
the pylon and wing tip fuel tanks and to operate the
thermal anti-icing system, the afterburner fuel pump,
the air-conditioning system, the canopy seal, and the
anti "G” suit. The main flow of air from the compres¬
sor then enters the eight combustion chambers where
fuel is sprayed under pressure and combustion occurs.
The hot combustion gases rotate a turbine wheel which
drives the compressor, both turbine wheel and com¬
pressor being mounted on the same shaft. From the
turbine wheel, the gases travel through the exhaust
cone and into the afterburner where additional fuel
may be injected and burned to create more thrust if
desired. The gases are then discharged from the tail¬
pipe. The afterburner tailpipe nozzle is equipped with
eyelids that open automatically during afterburning to
increase tailpipe diameter, thus allowing additional
thrust without excessive exhaust gas temperatures. The
afterburner eyelids, in addition to opening during
afterburning, will stay open during starting to prevent
high temperatures, and during rapid acceleration to de¬
crease acceleration time. Each engine at 100% rpm has
a rated thrust of 5600 pounds without afterburning and
7400 pounds with afterburning. Acceleration from idle
to 100% rpm requires approximately 12 seconds. For a
detailed discussion of the eyelids, see Eyelid Operation,
Section VII,
Each engine has one gear-type, constant displacement,
engine-driven fuel pump and one fuel control in¬
stalled in the accessory section. The maximum output
of each fuel pump is 26 gallons per minute. The en¬
gine-driven fuel pump incorporates two pumping ele¬
ments, Should one element fail, the other element will
maintain the required fuel pressure. Warning lights
(figure 1-9), located on the pilot's left console, will
indicate (on the ground only) that one of the pres¬
sure elements has failed. The fuel control automat¬
ically maintains the quantity of fuel supplied to
the engine within a range that will prevent "rich
blowout" during engine acceleration and "lean die-
out" during deceleration, and bypasses any fuel in
excess of that required by throttle setting, engine
speed, and altitude. For engine starting and controlled
acceleration during starting, the fuel is supplied to
the combustion chambers in a wide-angle spray for
ignition. This spray narrows its angle to distribute
the combustion more evenly throughout the chamber
as the engine accelerates. The change in spray charac¬
teristics is controlled within the nozzle by a spring-
loaded valve which opens another set of orifices in
the nozzle jet as fuel pressure builds up in the nozzle.
A centrifugal governor in the fuel control varies the
flow of fuel to the nozzles according to engine speed
and throttle position, (See figure 1-6.) Refer to Sec¬
tion VH for additional information on engine opera¬
tion.
Each of the two throttles (figure 1-7) on the pilot's
left console mechanically regulates an engine fuel con¬
trol. Markings on the throttle quadrant are CLOSED
and OPEN. Mechanical stops at the IDLE position pre¬
vent inadvertent retarding of the throttle below the
idle speed of the engines (49% to 51% rpm). The
throttles can be retarded past the idle stops by raising
the fingerlifts under the throttle knobs. This allows the
throttles to be placed at CLOSED, stopping fuel flow
to the engines. Each fingerlift connects to an after¬
burner demand switch that will start afterburning on
the corresponding engine when the throttle is between
the 90% and 100% rpm range. This is a mechanical
Throttles.
ENGINE FUEL CONTROL SYSTEM.
1-2
AtA/N DIFFERENCES TABLE
T.O. 1F-89H-1
Section I
««s
"°s
tn
w i-
W
U - a
P 5 rr
Figure 14*
1-3
Section I
T.O. 1F-89H-1
/™“~
POSITION LIGHT
RADAR OBSERVER'S EJECTION SEAT
RADIO AND ELECTRICAL EQUIPMENT
AUTOPILOT EQUIPMENT
EJECTION NOTIFICATION SWITCH
LOCALIZER ANTENNA
CIRCUIT BREAKER PANELS
PROBE &
i ome airplanes)
GUIDE AND HOOK
(Some airplanes)
RADIO COMPASS SENSE ANTENNA
1-4
Figure 1-4,
Changed 13 February 1959
T*0* 1F-89H-I
Section I
^ POSITION LIGHT (Each side)
1-5
Section I
T.O. 1F-89H-1
FUEL COMBUSTION CHAMBER
FUEL NOZZLE
AFTERBURNER EYELID CYLINDER
ACCESSORY GEAR DRIVE
I nipt Air Compressed Air
INNER CONE
0
AFTERBURNER EYELID ASSEMBLY
Combustion
0
Exhaust Gas
H-GB
Figure 1-5.
range on the throttle quadrant and is not dependent on
engine rpm. Afterburning is stopped by retarding the
throttle below the 90% position or by depressing the
fingerlift when the throttle is in the 90% to 100% rpm
range. The right throttle knob houses a press-to-talk
microphone button.
Throttle Friction Lever.
A throttle friction lever (figure 1-7) is provided on
the throttle quadrant outboard of the throttles. When
the lever is moved toward INCREASE or DECREASE,
resistance to throttle movement will increase or de¬
crease accordingly.
ENGINE COOLING AND AIR INDUCTION SYSTEM.
Engine cooling and induction air enters through an
air intake at the front of each engine. On the ground
and during takeoff, additional induction air is drawn
through four intake doors on the outboard side of the
engine forward door, and then through a door on the
engine transition duct. The combustion sections of the
engine compartment and the tailpipe are cooled by
ram air supplied through an airscoop on the lower
forward section of the engine's No. 3 and No. 4 doors.
Retractable screens in the engine air intakes normally
extend and retract with the landing gear, but under
certain conditions they can be operated during flight.
Vortex generators, in the form of two small air
directing vanes, are installed approximately 40 inches
forward of each engine air intake duct. The effect of
these vanes is to prevent the intermittent separation
of airflow through the engine transition duct and
the resultant noise and vibration which would occur
at high airspeed and low rpm.
ENGINE SCREENS.
Two retractable engine screens, one in each engine
air intake, provide a means for preventing foreign
matter from entering the engine intake ducts. The
engine screens normally extend and retract with the
landing gear; however, an engine screen switch pro¬
vides for screen extension and retraction during flight.
Engine Screen Switch.
The 2S-voIt d< engine screen switch (figure 4-4) on the
anti-icing control panel provides a means for extend¬
ing the engine screens during combat, or at other times
when there is danger of foreign matter entering the
engine intake ducts. The switch has two positions:
NORMAL and EMERG EXTEN. When the switch is
placed at NORMAL, the screens extend and retract
automatically with the landing gear. When the switch
is placed at EMER EXTEN, the screens extend; how¬
ever, if the anti-icing system operation is selected, the
screen control is overridden and the screens retract.
STARTING AND IGNITION SYSTEM.
Power for starting is supplied by 28-volt d-c external
power units connected to the power receptacles on
the right air intake duct. Only one engine can be
1-6
Changed 13 February 1959
From engine compressor
T.O. 1F-89H-1
Section i
Control System
ICfT CHCm (TYPICAL)
From a i rp la ne ‘s
fuel system
FILTER DE-ICE ALCOHOL TANK
ALCOHOL SHUTOFF VALVES
RIGHT
PRESSURE SWITCH
LANDING GEAR SAFETY RELAY
(Open when
airborne)
28-Y0LT DC BUS
COMPRESSOR!^_
AIR SHUTOFF
VALVE
yen during
afterbu ruing
operation)
HIGH-PRESSURE FUEL FILTER
HIGH-PRESSURE ^
FUEL FILTER
MAIN AFTERBURNER FUEL SHUTOFF VALVE
during afterburning operation ) |
28-VOLT D-C BUS 1
To fuel nozzles
To fuel nozzles
AFTERBURNER FUEL MANIFOLD
fuel nozzle (Typical)
From starling circuit
NORMAL FUEL FLOW
FUEL BYPASS
COMPRESSOR AIR
ALCOHOL - FILTER DE-ICE
ELECTRICAL ACTUATION
MECHANICAL ACTUATION
CHECK VALVE
SOLENOID VALVE
FUEL FLOWMETER INDICATOR
Figure 1-6 ,
1-7
To right engine
Section i
T.O, 1F-89H-1
Figure 1-7*
RADIO MIKE BUTTON
FRICTION LEVER
THROTTLES
FINGERLIFTS
started at a time* because actuating one starter-gener¬
ator breaks the d-c power circuit to the other engine's
starter and ignition system. When a starter switch is
actuated, the ignition system is energized and the
starter-generator cranks the engine. After the throttle
is opened and combustion is self-sustaining* the start¬
ing and ignition circuits automatically disconnect
when the electrical load drawn by the starter-generator
drops to 200 amperes; this should occur at an engine
speed of approximately 26% rpm. Then the starter-
generator functions as a 28-volt d-c generator. The
engine ignition system operates on 115-volt a-c power.
The essential bus of the single-phase inverter system
supplies current to the ignition transformers which,
in turn, send high voltage to the two igniter plugs in
each engine for both ground and air starting. The single-
phase inverter switch must be placed at NORMAL or
EMERGENCY before alternating current is available
for starting.
Starter and Ignition Switches.
Two starter and ignition switches (figure 1-12), one
for each engine, are located on the pilot's right ver¬
tical console. These switches have three positions;
START, NEUTRAL, and STOP. The switches are
spring-loaded to NEUTRAL. The switches, using
28-volt d-c power, control the electrical circuits to
the starter and to the 115-volt ignition system. When
a switch is at NEUTRAL, starter and ignition cir¬
cuits are open. Placing a starter switch momentarily
at START energizes the starter and completes the
circuit to the igniter plugs. When the load drawn by
the starter drops to 200 amperes, the starter and
ignition circuits automatically disconnect; this should
occur at an engine speed of approximately 26% rpm.
Placing the switch momentarily at STOP will de¬
energize the starter and ignition circuits. The starter
control circuit is interlocked with the fuel selector
switches, making it impossible to start an engine with
its fuel selector switch in the PUMPS OFF position.
This prevents loss of afterburner power during take-
off; however, this will have no effect on loss of after¬
burner power due to system malfunction or PUMPS
OFF fuel selector switch settings made after starting
the engines.
Altitude Start and Starter-Test Switches.
Two altitude start and starter-test switches (figure
1-17), one for each engine, are located on the aft
miscellaneous control panel above the pilot's left con¬
sole. These switches are for ignition during air
starts and for turning the engine over by the starter
without ignition. The switches have three positions;
1-8
T.O. 1F-B9H-1
Section I
TEST, NEUTRAL, and START; they are spring-
loaded to NEUTRAL. The switches, using 28-volt d-c
power, control separate electrical circuits to the 115-
volt a*c ignition system and the 28-volt d-c starter.
When a switch is at NEUTRAL, starting and ignition
circuits are open. When an air start is required, plac¬
ing the switch momentarily at START will supply
ignition to the windmilling engine for 120 seconds
through a time-delay unit. When the switch is held at
TEST (for ground operation only), the starter will
turn the engine over without ignition.
Starting Power Switch.
A guarded switch (figure 1-12) with two positions,
EMERGENCY and NORMAL, is located on the pilot’s
right vertical console. This switch connects the 28-volt
d-c primary bus to the starter bus for emergency starting
when limited external power is available. When only
one 28-volt d-c external power source is available (of at
least 1000-amp rating), the one lead may be plugged into
the lower d-c receptacle (with the battery switch at
OFF) and the starting power switch placed at EMER¬
GENCY.
Note
If only one external power unit is available
for starting the airplane, it must be of at
least 1000-amp capacity.
The engine then can be started with the starter
switches. When the starting power switch is at NOR¬
MAL, the starter bus is disconnected from the 28-volt
d-c primary bus.
CAUTION j:
For emergency starts, the 28-volt d< genera¬
tor switches must be at OFF. This is to pre¬
vent the left generator from overloading dur¬
ing the right engine start.
Note
This airplane cannot be started on the bat¬
tery. External power is required.
EXHAUST GAS TEMPERATURE GAGES.
Two exhaust gas temperature gages (figure 1-8), in¬
dicating exhaust temperature in degrees centigrade,
are located on the pilot’s instrument panel. The gages
operate from thermocouples located in each engine
exhaust cone and are independent of the airplane’s elec¬
trical system. There is no direct control for regulating
the exhaust temperatures by the pilot; however, limited
control for these temperatures can be indirectly achieved
by changing the throttle settings. See Section VII for
a discussion of exhaust gas temperature versus runway
temperature.
TACHOMETERS.
Two tachometers (figure 1-8), indicating engine speed
in % rpm, are located on the pilot’s instrument panel.
A tachometer generator is installed in the accessory
section of each engine. The electrical power and fre¬
quency it produces for tachometer readings is propor¬
tional to engine rpm (100% engine speed is 8000 rpm).
OIL PRESSURE GAGES.
Two oil pressure gages (figure 1-8), one for each
engine, are located on the pilot's instrument panel and
indicate oil pressure in pounds per square inch. The
gages are operated by 115-volt ac from the single¬
phase inverter essential bus.
FUEL FLOWMETER INDICATORS.
Two fuel flowmeter indicators (figure 1-8), one for
each engine, are located on the pilot’s instrument
panel. The indicators show rate of flow in pounds
per hour and use both 28-volt dc and 115-volt ac.
Note
When the afterburner is operating, a rise in
fuel flow will be experienced; however, the
fuel flowmeter indicators do not indicate fuel
consumed by the afterburners.
I; CAUTION !j
The fuel flowmeter indicators are inaccurate
for high rates of fuel flow. However, in the
cruising range (3000 to 5000 pounds per hour),
the indicators may be relied upon for cruise
control.
ENGINE-DRIVEN FUEL PUMP FAILURE WARNING
LIGHTS.
Two 28-volt d-c fuel pump failure warning lights (fig¬
ure 1-9), one for each engine, are located on the pilot’s
left console. The lights are provided to warn the pilot
that one of the two elements of the engine-driven fuel
pumps is inoperative. The lights are controlled by a
pressure switch connected to the two pumping elements.
If the fuel pressure drops at the outlet of one element,
the switch closes and turns on the light. The lights will
indicate an element failure during ground operation
only. A switch on the left main landing gear prevents
operation when the weight of the airplane is removed
from the landing gear.
AFTERBURNER SYSTEM.
Each engine has an afterburner which can be used to
increase thrust when needed. The afterburner is a
part of the tailpipe. As the gases travel through the
exhaust cone and into the afterburner section, more
fuel can be injected and burned if additional thrust
is desired.
1-9
Section I
T.O. 1F-89H-1
T.O. IF-89H-1
Section I
h
z
Figure 1-9*
1-11
Section I
T.O, 1F-09H-T
Normal fuel sequencing must be used to main¬
tain afterburning. For further explanation
refer to Section YIL
Afterburning is best initiated from a stabilized
full-throttle condition. A speed-sensing switch pre¬
vents afterburner ignition when engine speed is be¬
low 87*5% rpm* The afterburner fuel control system
(figure 1-6) consists of a centrifugal-type fuel pump
which is driven by an air turbine powered by air bled
from the engine compressor* This pump supplies fuel
to an afterburner fuel regulator* The fuel regulator,
controlled by the difference in pressure between the
inlet and the outlet of the engine compressor, auto¬
matically meters a continuous flow of fuel to the
afterburner. When afterburning is initiated (by lift¬
ing the fingerlifts on the throttles), the following
operations take place in the automatic control system
LANDING AND TAXI
LIGHT
PYLON TANKS JETTISON BUTTON
CANOPY SEAL
VALVE LOCK
CANOPY SEAL VALVE BUTTON
LANDING GEAR
EMERGENCY OVERRIDE LEVER
LANDING GEAR LEVER
LANDING GEAR
EMERGENCY RELEASE HANDLE
pilots ten
VtRTKAl COHSOU
H-l 3C
1-12
Figure I-JO,
PULL
TO CAGE |
SIGHT
Section 1
TO. 1F-89H-1
0 If both single phase inverters fail below
10,000 feet, or if the afterburner a-c control
circuit breaker pops out, the afterburner and
afterburner circuit will be inoperative. When
this occurs, the throttle-actuated eyelid
switches will cause the eyelids to open (with¬
out regard to afterburner operation) when
the throttles are advanced to OPEN, resulting
in very low tailpipe temperatures and ex¬
treme loss of thrust. If both inverters fail
below 10,000 feet while in afterburning, after¬
burner operation will be unaffected. How¬
ever, if the afterburners are shut down by de¬
pressing the throttle fingerlifts, the eyelids
will remain open. The eyelids must be closed
by moving the afterburner control circuit
breakers to OFF or by retarding the throttles
STARTER-IGNITION SWITCHES
—® EMERGENCY SIGNAL
BUTTON AND LIGHT
STARTING POWER SWITCH
BATTERY SWITCH
FIRE EXTINGUISHING
-—® CONTROL PANEL
CANOPY JETTISON HANDLE
^ PILOTS RADAR
INDICATOR KNOB
ALTERNATOR CONTROL PANEL
MOT'S RIGHT
VERTICAL CONSOLE
HT5G
T-14
Figure 1- 12.
Section i
T.O. 1F-89H-1
Figure V-T3
Section [
T*Q* 1F-89H-1
to approximate 90% position. Eyelid closure
will be apparent by an immediate increase in
thrust and a return to normal tailpipe tem¬
perature. Only military thrust will be avail¬
able for the duration of the flight.
@ If both single-phase inverters fail while in
afterburning above 10,000 feet, afterburning
will be unaffected because the holding relay
in the afterburner control box keeps the eye¬
lids open; however, once afterburning is shut
off, it cannot be reinitiated* If both single-
phase inverters fail above 10,000 feet, after¬
burning cannot be initiated and eyelids will
remain in the closed position, because the
altitude switch breaks the d-c operating cir¬
cuit, allowing the "fail safe” eyelid control
valve to keep the eyelids closed*
© Use the ALL TANKS fuel selector position
for afterburner takeoffs, as this position af¬
fords a greater margin of fuel pressure for
maintaining afterburner operation than the
WING TANKS selector position because less
flow resistance exists in the distribution lines
from the main tanks*
AFTERBURNER DEMAND SWITCHES,
Two afterburner demand switches control afterburner
operation* Each switch is connected by mechanical
linkage to a fingerlift on each throttle (figure 1-7).
The switches use 28-volt dc to control the electrical
circuits in the automatic afterburner system* After¬
burning is initiated by lifting the fingerlift when the
throttle is in the 90% to 100% rpm range* A speed¬
sensing switch for each engine prevents afterburner
ignition when engine speed is below 87,5% rpm*
When a fingerlift is raised and engine speed is above
87*5% rpm, the valve controlling compressor air to
the turbine-driven afterburner fuel pump opens, the
main afterburner fuel shutoff valve opens, and hot-
streak ignition occurs. After the fuel ignites, the eye¬
lids open, and afterburning continues. Both afterburners
may be ignited at the same time during scrambles or in
an emergency* However, individual ignition is recom¬
mended to check ignition of each burner* After¬
burning is stopped by depressing the fingerlift when
the throttle is in the 90% to 100% rpm range or by
retarding the throttle below the 90% rpm position.
Either action will turn off the afterburner demand
switch* The fingerlift does this by direct mechanical
linkage, and retarding the throttle does it by means
of a cam arrangement in the throttle quadrant. If
afterburning is stopped by retarding the throttle, the
fingerlift will be lowered to the down position.
Stopping afterburning by either method returns all
units of the automatic control system to a non afterburn¬
ing condition and restores normal engine operation. If
the afterburner flames out, the automatic control will
shut down the afterburner. The afterburner will not
reignite until the fingerlift is depressed and then raised
again while the throttle is in the 90% to 100% rpm
range and engine speed is above 87.5% rpm.
AFTERBURNER CONTROL SWITCHES.
Two toggle-type afterburner control switches (circuit
breakers figures 1-9 and 1-13), one on each 2S-volt d-c
circuit breaker panel, are used to deenergize the after¬
burner control circuits during ground operation or dur¬
ing afterburner malfunction in flight. Each switch has
a placarded ON and an unmarked OFF position* When
placed at ON, the afterburner control circuits are ener¬
gized; when placed at OFF (unmarked), the circuits
are deenergized. If more than 15 minutes are to
elapse between supplying power to the 28-volt d-c
bus and starting or operating engines above idle rpm,
place the afterburner control switches at OFF (un¬
marked) and leave them OFF until just before starting
engines. This will deenergize the altitude idle bleed
and eyelid actuator solenoids, thus preventing them
from being damaged by overheating. Two a-c push-
pull circuit breakers, one for each afterburner, are lo¬
cated on the a-c circuit breaker panel on the pilot's
right vertical console (figures 1-13 and 1-25), When the
circuit breaker is set (pushed IN), inverter power acti¬
vates the speed-sensing switch which in turn controls
the eyelids at 87^% rpm and above (open during
afterburner operation, closed during nonafterburner
operation)* Below 87V^% rpm, the eyelids are con¬
trolled by the lower microswitch located on the engine
throttle quadrant. If the circuit breaker is deactivated
(pulled or pops out) the eyelids will go to the open
position when the upper engine throttle quadrant
microswitch is actuated (approximately 90% rpm)
which will be denoted by a decrease in engine gas
temperature and a loss of thrust. See Afterburner, this
section*
AFTERBURNER WARNING LIGHTS.
Two afterburner warning lights (figure 1-8), located
on the pilot's instrument panel, provide a visual check
of eyelid position during afterburner operation* When
the engines are being operated at 87.5% rpm or above
and afterburning is selected, the two warning lights
will come on* These lights will stay on {usually I to 5
seconds) until afterburner eyelids open, at which time
the lights will go off, indicating normal after¬
burner operation* If the warning lights fail to go off
(indicating eyelids closed), afterburning should be
discontinued. If an afterburner flanieout occurs, the
warning light will not come on until the eyelids close.
If the eyelids do not close, the flameout will be indi¬
cated by the increase in rpm and a sudden drop in
tailpipe temperature on that engine* The afterburner
warning lights are inoperative on some airplanes.
1-16
T.O. 1F-89H-1
Section 1
OIL SUPPLY SYSTEM.
Each engine has an independent dry sump, full scav¬
enge oil supply system. See figure 1-1 4 for oil quantity
data. Oil is gravity fed from the tank, mounted on the
outboard side of the engine, to the main engine-driven
pump. The main pump distributes the oil under pres¬
sure through a filter to the accessory gears and engine
bearings. The scavenge side of this same pump returns
oil from the accessory and forward engine bearing to
the oil tank, A mid frame scavenge pump scavenges
oil from the mid, damper, and aft bearings, and re¬
turns it through a heat exchanger to the oil tank. The
heat exchanger uses fuel flow to cool the scavenged
oil. The operation of this system is entirely automatic.
See figure 1-45 for oil specification and grade.
FUEL SUPPLY SYSTEM,
The airplane has two independent fuel supply systems*
left and right, with interconnecting lines and valves for
crossfeeding (figure 1-16), Each system has a main
fuselage tank, two multiceiled wing tanks, a perma¬
nently installed tip tank, and a jettisonable pylon tank.
The right system main tank is in the nose section; the
left system main tank is in the aft fuselage. For fuel
quantity data, see figure 1-15, During normal opera¬
tion, fuel is pumped to the engines from the main
tanks which are automatically replenished from the
wing tanks. As wing tank fuel level is lowered, fuel
from the pylon and tip tanks flows simultaneously into
the wing tanks under air pressure from the engine
compressors. Fluid level actuated valves within the
wing tanks close when the tanks become full to pre¬
vent overfilling and pressurization. Pressurization of
pylon and tip tanks is automatically regulated at ap¬
proximately 6 psi and transfer of fuel from these tanks
will continue until the pylon and tip tank fuel supply is
exhausted or jettisoned. When a tip tank empties, a
fluid level actuated shutoff valve within the tank
closes the fuel line from the empty tank to the wing
tanks. When a pylon tank empties, a float switch is
acruated, causing a solenoid valve in the pylon air
pressurization line to close. After the wing tanks be¬
come empty, the main tanks continue to supply fuel
to the engines. When the main tank fuel level is low¬
ered to the 100-gallon (650-pound) level, a low-level
warning light on rhe pilot's instrument panel will
glow red. For all normal operation, fuel flow sequence
is completely automatic; however, wing tanks may be
selected and fuel will be pumped directly from wing
tanks to the engine. Cross feed operation permits both
engines to operate from either fuel system, or permits
single-engine operation from either or both fuel sys¬
tems. Fuel for afterburning is pumped from the main
Figure J-14 .
fuel line to a turbine-driven pump on each engine,
through the afterburner fuel regulators, and then to
the afterburners, (For fuel specification and grade, see
figure 1-45.)
Booster Pumps,
Each of the two fuel systems has four 28-volt d-c
booster pumps, one in each of the wing tanks and two
in each main tank. During normal operation all
booster pumps operate continuously. The pumps are
designed for sustained operation wet or dry, and
therefore may operate in an empty tank.
Low Pressure FueB Filter De-Icing System,
A low pressure fuel filter de-icing system is provided
for the engines. Alcohol is injected at the pilot's dis¬
cretion into the low-pressure filters to dissolve any ice
accumulation in the filters or engine fuel controls.
(For further description and operating procedure for
this system, refer to Section IV. For alcohol specifi¬
cation, see figure 1-45,)
M7
Section [
T.O. 1F-89H-1
US GALLONS 1
’POUNDS f
(EACH TANK)
PYLON OUTBD iNBD RIGHT LEFT
(2) WING (2 > WING <2) MAIN MAIN
GAL
LB
GAL
LB
GAL
LB
GAL
LB
GAL
LB
GAL
:z
i
r
30*
j
i
1989
300
| 1950
246
1599
104
676
261
i
i
1696
196 !
1274
612
i
;
3978
600
j 3900
492
3198
208
1352
261
i
i
1696
196 !
i
1274
308
r
i
i
2002
301
| 1956
252
1633
106
689
263
i
!
1709
200 j
1300
616
t
i
4004
602
j 39T2
504 ;
3276
212
1378
263
1
3
1
1709
200 |
i
1300
pylon lank fuel 2369 gal funs (15,398 lh*)
Without pylon tank fuel 1769 gallons (11,498 lb*)
DATA AS OF / June 1955
DATA BASIS; Calibration
H47C
figure I-F5.
Pylon Tank Jettison System,
The pylon tanks may be jettisoned electrically or, in an
emergency, released manually. The ejection system
in each tank pylon includes an ejector mechanism,
consisting of an electrically ignited propellant charge.
When the pylon tanks are ejected, 28-volt d-c power
ignites the propellant charge which releases the attach¬
ing hooks and actuates an ejection piston which forci¬
bly ejects the tanks clear of the airplane. When the
tanks are manually released by pulling the external
stores emergency release handle, they fall by gravity
alone. Both tanks are jettisoned simultaneously.
If pylon tanks are manually released, minor
damage to the airplane may occur.
MS
Tip Tank Fuel Dump System.
Each tip tank has a 28-volt d-c motor-driven dump
valve located in the tip tank tailcone. When these
valves are opened, fuel is forced overboard under nor*
mal tip tank air pressure through an outlet in the tail¬
cone at a rate that will normally empty a full tip tank
in approximately 90 seconds. The valves are held open
for approximately 2 minutes by a time-delay relay.
Tip tank fuel will not be completely dumped during de¬
celerations or dives; however, a new dumping cycle may
be initiated if required.
Note
# Tip tank fuel cannot be dumped while the
weight of the airplane is on the wheels, be¬
cause the oleo strut ground safety switch
breaks the tip tank fuel dump electrical circuit.
• The fuel gage selector switch should be placed
at the TIP position prior to and during dump¬
ing of tip fuel. This will enable the pilot to
T,0. 1F-89H-I
Section I
determine if the fuel in both tip tanks has
been dumped and whether or not an un¬
balanced tip tank fuel condition exists*
Single-Point Fueling System,
For description and operation of the single point fuel¬
ing system, refer to Section IV.
FUEt SELECTOR SWITCHES.
Two rotary 28-yolt d-c selector switches (figure 1-18),
one for each system, are located on the fuel control
panel. Each switch has ALL TANKS, WING TANKS,
and PUMPS OFF positions. When a selector switch is
at ALL TANKS, all related booster pumps operate
continuously and fuel sequencing is automatic: pylon
and tip tanks feed the wing tanks, wing tanks feed
the main tank, and the main tank feeds the engine.
When a selector switch is at WING TANKS, only
the wing tank booster pumps in that system operate
and fuel is routed directly from wing tanks to the
engine; however, pylon and tip tanks will continue
to replenish the wing tanks. When a selector switch
is at PUMPS OFF, all booster pumps in that system
are shut down. The starter control circuit is inter¬
locked with the fuel selector switches, making it im¬
possible to start an engine with Its fuel selector in the
PUMPS OFF position. This modification prevents loss
of afterburner power during takeoff but has qo effect
on loss of afterburner power due to system malfunc¬
tion or PUMPS OFF power switch settings made after
starting engines.
Note
Placing the selector switch at PUMPS OFF
does not close the firewall fuel shutoff valve.
This valve will close when the throttle is
moved to the closed position or when the
engine fire selector switch is actuated.
After positioning the fuel selector switch at
any position, allow at least 3 seconds to elapse
before selecting another position. This will
preclude any possibility of the affected fuel
system motor valves being reversed in mid¬
cycle, thus shortening valve life.
| CAUTION
system. Unbalanced lateral fuel loading (wing heavi¬
ness ) may be corrected by feeding both engines from
the system having more fuel. To balance fuel load,
the crossfeed switch is placed at OPEN and the fuel se¬
lector switch for the system with less fuel is placed at
PUMPS OFF. When fuel load is balanced, as indicated
by lateral trim or fuel quantity gages, the selector
switch is returned to ALL TANKS and the crossfeed
switch to CLOSED.
ENGINE FIRE SELECTOR SWITCHES.
Two guarded 28-volt d-c engine fire selector switches
(figure 1-39), one for each engine, are located on the
pilot’s right vertical console. Lifting the guard and
placing either switch in the UP position arms the
fire extinguishing agent discharge switch and closes
those fuel shutoff valves which isolate the related
engine from its fuel supply.
THROTTLE-ACTUATED FUEL SHUTOFF SWITCHES,
Two 28-volt d-c throttle-actuated fuel shutoff switches,
one for each engine, are actuated when the throttles
are moved to the closed position. Actuation of these
switches closes the firewall fuel shutoff valves.
Note
If the right engine fuel selector switch is at
WING TANKS, the related throttle-actuated
fuel shutoff switch will not isolate the en¬
gine from its fuel supply.
PYLON TANKS JETTISON BUTTON.
A 28-volt d-c pushbutton (figure 1-10) marked PY¬
LON TANKS JETTISON is located on the pilot’s
left vertical console. When the button is pressed, both
pylon tanks are ejected simultaneously,
EXTERNAL STORES EMERGENCY RELEASE HANDLE.
An external stores emergency release handle (figure 1-9)
is located on the pilot’s left console. This emergency
release handle is linked by cables and bellcranks to the
bomb shackle release in each pylon. When the handle
is pulled out approximately 7 inches with a force of
approximately 30 pounds, both right and left bomb
shackles will be tripped simultaneously and both pylon
tanks will drop by gravity.
j; CAUTION j;
CROSSFEED SWITCH.
A 28-volt d-c crossfeed switch (figure 1-18), located
on the fuel control panel, has OPEN and CLOSED
positions. When the crossfeed switch is at OPEN, the
main fuel lines of both systems are interconnected;
both fuel systems may be used to operate one engine
or both engines may be operated from either fuel
If pylon tanks are manually released, minor
damage to the airplane may result.
TIP TANK FUEL DUMP BUTTON.
A 28-volt d-c pushbutton (figure 1-18) marked PRESS
TO DUMP TIP TANK is located on the fuel control
panel. When momentarily pressed, this switch operates
1-19
Section I
T.O. 1F-89H-I
From engine compressor
ARD TANK
LEF
MAIN .’A!
(LOCA I il)
PYLON TANK
o engine
fuel
control
system
CLOSED
BOOSTER PUMP
FLOAT SWITCH
SINGLE-POINT FUELING ONLY
BREAKAWAY CONNECTION
SOLENOID VALVE
{SPRING LOADED TO CLOSED)
COMPRESSOR AIR
ELECTRICAL ACTUATION
CROSSFEED SWlTl
FUEL LEVEL SENSING UNIT
CHECK VALVE
FUEL LEVEL ACTUATED
SHUTOFF VALVE
PRESSURE-VACUUM
RELIEF VALVE
i
*
1-20
Figure 7-16.
T.O. 1F-89H-1
Section I
—1
Left electrical circuits same as shown
on right side {except for aft C-G warning
light system which exists only on
right side)* For normcd positions of valves
and controls during various fuel flows*,
see figure 7-1*
L.H. MAIN AFT C.G, R.H. MAIN
LOW LEVEL WARN WARNING LOW LEVEL WARN
INBD
MAIN. • _OUTBD
H-19(2>C
1-21
Section 1
T.O. 1F-89H-1
I'iloi'x cockpit-luff side
PILOT’S MISCELLANEOUS CONTROL PANEL
Figure I-17*
a time-delay relay -which opens right and left tip tank
dump valves for approximately 2 minutes* A full tip
tank will normally dump in approximately 90 seconds;
however, during dives or decelerations, all tip tank fuel
will not be dumped. A new dumping cycle may be initi¬
ated if required.
FUEL SYSTEM WARNING LIGHTS*
One aft eg and two low-level warning lights, all oper¬
ating on 28-voit dc, are located on the pilots instru¬
ment panel immediately above the fuel quantity indi¬
cators* Each low-level warning light is operated by a
Do not dump tip tank fuel when firing
rockets or missiles because of the fire hazard.
FUEL QUANTITY GAGES A NO SELECTOR SWITCH*
Two 115-volt a-c fuel quantity gages and a five-posi¬
tion 28-volt d-c rotary tank gage selector switch (fig¬
ure 1-19), located on the pilot's instrument panel, en¬
able the pilot to read total fuel quantity or individual
tank quantities* Positions of the gage selector switch
are ALL, MAIN, INBD, OUTBD, and TIP* (Pylon tank
fuel is not included in the fuel quantity gage system*)
When the gage selector switch is placed at ALT. , total
fuel in each system is indicated on the respective quan¬
tity gage; when the switch is placed at any individual se¬
lection, only the amount of fuel in the selected tank is
indicated. Capacitance-type fuel probes are located in
each tank and vary an electrical signal in proportion to
fuel level; the resultant signal changes are reflected on
quantity gages calibrated in pounds of fuel* To deter¬
mine that the gages are operating, a 28-volt d-c press-to -
test switch, common to both gages, is provided on the
panel between the two quantity gages* Pressing this
switch causes the needles of the gages to swing to off-
scale positions; releasing the switch causes the needles
to return to their original positions, thus indicating
that the fuel quantity system is functioning*
{Riot's left roust tie
FUEL CONTROL PANEL
*nc
Figure 1 - 18 .
1-22
T.O. 1F-89H-1
Section I
/MAIN l OUT BO 9
W ' '
'ALL, TIP
L.H. MAIN
LOW LEVEL WARN
AFT CjG.
WARNING
Quantity 4
I 1(H * IODO 5 -
R.H. MAIN
LOW
r 1 n ° /\
2 FUEL ^ V
QUANTITY
t LiU * 1000 5 - \
FUEL QUANTITY
SEL ECTOR SWITCH
I [PILOT’S CHECK L
1 F 89 AIRPLANE
rFEB <955
LIST
E V
(jit 1/-
FUEL QUANTITY
TFKT
Pilot s instrument panel
roti aoANmy caccs
H-22C
Figure 1-19,
float switch in each related main tank and will come
on when main tank fuel is lowered to the 100-gallon
(650-pound) leveL The aft eg warning light is oper¬
ated by, and in series with, two float switches, one
located near the full level of the right main tank and
one located near the empty level of the right tip tank.
The aft eg warning light will come on when the main
tank fuel level is lowered 50 gallons (325 pounds)
from full with any fuel, above residual, remaining in
the right tip tank* When the aft eg warning light
comes on, airspeed must be reduced to Mach 0*65 or be¬
low* For discussion on center of gravity limitations, refer
to Section V; for corrective action for aft eg warning,
see Section HI*
ELECTRICAL POWER SUPPLY SYSTEMS*
One direct-current system and three alternating-cur¬
rent systems supply the electrical power* The 28-volt
d-c system obtains power from three engine-driven
generators, one on the left engine and two on the
right engine* A 24-volt, 36 ampere-hour storage bat¬
tery in the forward fuselage section serves as standby for
emergency d-c circuits* The d-c generator on the left
engine and one of the d-c generators on the right engine
also function as starters* Full generator output is
reached at 35% engine rpm* Alternating current is sup¬
plied by a constant frequency 115-volt a-c single-phase
inverter system, a constant frequency 115-voh a-c three-
phase inverter system, and a variable frequency 115/200-
volt a-c three-phase alternator system. All inverters, two
for each system, are powered by the primary 28-volt
d-c bus. The alternator is engine-driven and is located
on the left engine. External a-c power is required for
ground operation and starting. External power recep¬
tacles for the 28-volt d-c system and 115/200-volt a-c
alternator system are on the right engine air intake
duct.
1-23
Section I
T.Q, 1F-89H-I
ELECTRICAL SYSTEM LOAD DISTRIBUTION TABLE
POWER SOURCE LOST
INVERTERS:
1. POWER
a, 115-volt AC
single-phase
2500-VA (main)
115-volt A C
single-phase
2500-VA (spare)
EQUIPMENT LOST
AFTERBURNER SPEED-SENSING
SWITCH
AUTOPILOT
AUTOSYW INSTRUMENTS
CABIN TEMPERATURE CONTROL
ENGINE IGNITION
FUEL QUANTITY GAGE SYSTEM
GLIDE SLOP! RECEIVER
SIDESLIP STABILITY AUGMENTER
VHF NAVIGATION RECEIVER
WINDSHIELD DE-ICE AND DEFOG
CONTROLLER
FLIGHT COMPUTER
DIRECTIONAL INDICATOR
(SLAVED)
FIRE CONTROL SYSTEM
EQUIPMENT PICKED UP
AUTOMATICALLY
FLIGHT COMPUTER
DIRECTIONAL INDICATOR
NONE
EQUIPMENT PICKED UP
MANUALLY
By manually selecting
emergency operation, the
&jiare inverter will supply
power to all equipment
normally powered by the
main inverter *
NONE
EQUIPMENT LOST
PERMANENTLY
Power to the Fire Control
System will be cut off when
power from the spare
inverter is shifted to the
Essential bus upon select¬
ing emergency operation*
FIRE CONTROL SYSTEM
NONE
Only one inverter 7 main
or spare f operates at a
time; if one fails , select
the other,.
NONE
FIRE CONTROL SYSTEM
FUEL VENT DE4CE HEATERS
MISSILE HEATERS
NADAR HEATER
RADAR
RADOME ANTMCING FLUID
HEATER
WINDSHIELD DE-ICE HEATER
NONE
. INSTRUMENT
115-volt AC
three-phase
500-VA
a. main
i>. spare
ATTITUDE INDICATOR
ALTERNATOR:
200/115-volt A C
three-phase
FIRE CONTROL SYSTEM
FUEL VENT DE-ICE HEATERS
MISSILE HEATERS
NADAR HEATER
RADAR
R A DOME ANTMCING FLUID
HEATER
WINDSHIELD DE-ICE HEATER
FIGHTER IDENTIFICATION
SYSTEM
WINDSHIELD DEFOG
FIGHTER IDENTIFICATION
SYSTEM
WINDSHIELD DEFOG
GENERATORS:
BATTERY:
24-volt, 36 ampere-
hour storage
NONE
(The battery serves as
standby for D C circuit
during flight.)
28-volt D C genera¬
tors (One on left
engine and two on
right engine)
NONE
// one generator finis*
the remaining tico ivill
carry the toad ,
NONE
NONE
Figure 1-20.
T-24
T.O. T F-89H-T
Section 1
Electrically Operated Equipment.
For complete reference of power distribution to elec¬
trically operated equipment, see figure 1-21.
External Power System.
Two 28-volt d-c and one 115/200-volt three-phase a-c
external power receptacles provide a means of starting
the engines and operating all electrical equipment from
external power. The three external power receptacles are
located on the right engine air intake duct. The top
receptacle is for 28-volt d-c starting power only. The
center receptacle is for external power to the 28-volt
d-c distribution bus. The lower receptacle is for ex¬
ternal 115/200-volt three-phase 400-cycle power to
the alternator and inverter buses. D-c loads are auto¬
matically assumed by the external power sources. To
transfer a-c loads to external a-c power, the alternator
breaker control switch must be momentarily placed at
TRIP, the exciter control switch must be momentarily
placed at CLOSE, the external power switch must he
momentarily placed at CLOSE, then the single-phase
inverter switch must be placed at NORMAL and the
three-phase inverter switch placed at MAIN. External
115/200-volt three-phase a-c power is then connected
to the alternator distribution bus; single-phase 115-
volt a-c power is connected to the single-phase essen¬
tial and secondary buses; and three-phase 115-volt a-c
power is connected to the three-phase a-c bus. To
transfer a-c loads from the external a-c power to the
airplane's a-c power (after engines are running and
two or more d-c generators and the alternator are
operating), either the external power switch must be
placed momentarily at TRIP or the alternator breaker
switch placed momentarily at CLOSE. The airplane's
a-c power system will then be in normal operation.
Mote
® A-C loads will automatically transfer from ex¬
ternal a-c power to the airplane's a-c power
(inverter, alternator, and generator switches
set for normal operation) when external a-c
power is removed from the airplane. When
a-c loads are being carried by an external
power supply, the alternator circuit is open
and the single-phase and three-phase invert¬
ers will not operate.
© If the three-phase inverter switch is placed
at SPARE or the single-phase inverter switch
at EMERGENCY, the external a-c power
source will be automatically disconnected from
the airplane.
!J CAUTION ;;
Three-phase a-c external power must be used
with this airplane. Single-phase a-c power will
damage airplane equipment.
28-VOLT D-C SYSTEM.
The 28-volt d-c system obtains power from three en¬
gine-driven generators, one on the left engine and two
on the right engine. The d-c generator on the left
engine and one of the d-c generators on the right en¬
gine also function as starters. The two starter-genera¬
tors crank the engines until the electrical load drops to
about 200 amperes (approximately 26% rpm) and then
all three generators cut in after engine speed reaches
28% rpm. Three bus bars provide for distribution of di¬
rect current; a battery bus, a primary bus, and a second¬
ary bos. When the engines are being cranked, reverse-
current relays disconnect the d-c generators from all
but the starter bus. When the engines are operating,
the three d-c generators supply both the primary bus
and the secondary bus, and the two bus bars are in¬
terconnected by a bus-tie relay. Failure of any two
generators will separate the two buses, and the remain¬
ing d-c generator will supply power to the primary
bus only. A 24-volt 36 ampere-hour storage battery is
connected in series to the main 28-vok d-c bus through
the battery relay. If all three 28-volt d-c generators
fail, the battery will operate emergency 28-volt d-c
equipment for a limited time. If an emergency start is
necessary, with one 28-volt d-c external power source
available, an emergency bus-tie relay (through the
starting power switch) connects the primary 28-volt
d-c bus (energized by plugging external power into
the lowest d-c receptacle) to the starter bus. With the
exception of the battery switch on the pilot's right
vertical console, all controls and indicators for the
28-voit d-c system are on the pilot's right console.
Battery Switch.
The battery switch (figure 1-12), located on the pilot's
right vertical console, connects the battery bus with
the 28-volt d-c primary bus and has ON and OFF
positions. When the switch is at ON, the battery bus
is connected to the 28-volt d-c primary bus. Whenever
the 28-volt d-c system is operating and the battery
switch is at ON, the battery is being charged. When
the switch is at OFF, the circuit connecting the battery
bus to the primary bus is broken.
MOTION ;;
The battery switch must not be at ON when
the external 28-vok d-c starting power supply
is being used to start the engines, as damage to
the battery will result.
28‘-Volf D-C Voltage Regulator Rheostats.
Three voltage regulator rheostats (figure 1-13), one
for each 28-volt d-c generator, are located under a
hinged cover next to the 28-volt d-c indicator and
control panel on the pilot's right console. The 28-volt
1-25
Section I
T.O. 1F-89K-1
INCH
RIGHT ENGINE
NO. 7
STARTER-GENERATOR
I NCR
DEC
RIGHT ENGINE
NO. 2
GENERATOR
BUS-TIE RELAY
Energized when two
or more d< generators
are operating.
EXTERNAL POWER UNIT
FROM 28-VOLT DC
GENERATOR CONTROL CIRCUITS
a
N-230JB
BATTERY BUS
Energized by battery at all
times, by 28 -volt d-c bus
when battery switch is ON.
ARMAMENT JETTISON
CANOPY OPERATION
EMERGENCY RIGHT CONTROL PUMP
PYLON TANK JETTISON
RANGE LIGHTS. PILOT'S SCOPE
SNAKE LIGHT
28 VOLT D-C PRIMARY BUS
Energized by generators or
external power units;
connected to battery bus
when battery switch is ON .
AFTER BURNER CONTROL
ALTERNATOR ANO MAIN SINGLE-PHASE
INVERTER CONTROL
COCKPIT, LANDING-TAXI LIGHTS
AND POSITION LIGHTS
COMMAND RADIO
M GENERATOR CONTROLS
ENGINE CONTROL
ENGINE SCREEN COMPRESSOR
EXTERNAL A*C POWER CONTROL
EIRE CONTROL SYSTEM
EIRE DETECTOR AND EXTINGUISHER
RIGHT COMPUTER AND REMOTE COMPASS
FREE AIR TEMPERATURE INDICATION
FUEL FILTER DEICE CONTROL
FUEL SYSTEM AND CONTROLS
GLIDE SLOPE ANO OMNIRANGE
HYDRAULIC PRESS. CUTOFF
HYDRAULIC PUMP, LEFT SUPPLEMENTAL
HYDRAULIC RESERVOIR TEMPERATURE
CONTROL
I FT ANO FIS SYSTEM
INSTRUMENT PANEL VIBRATORS
INTERPHONE
INVERTERS
LANDING GEAR INDICATION
LANDING GEAR SAFETY RELAYS
LANDING GEAR WARNING
MARKER BEACON
NOSE WHEEL STEERING
OVERHEAT WARNING SYSTEM
OXYGEN WARNING
PHOT TUBE HEATERS
RADAR BLOWERS
RADIO COMPASS
RADOME ANTMCE CONTROL
STARTER-IGNITION CONTROL
THREE-PHASE INVERTER CONTROL
TRIM CONTROL
TURN AND SUP INDICATOR
VERTICAL GYRO AUTOPILOT
WINDSHIELD ANTI-ICING
WINDSHIELD WIPERS
SECONDARY BUS
RADAR COMPRESSOR
RADAR POWER CONTROL
SPARE SINGLE- PHASE INVERTER CONTROL
T-26
Figure 1-21.
T.O. 1F-89H-1
Section I
NORMAL
EM*ENCY'
SINGLE-PHASE
SECONDARY B( S
Energized by spare inverter
when switch is at NORMAL.
VOLTMETER AND
SELECTOR SWITCH
l f I
PHASE CONVESUR
(Sin#U~pkuse to
Three-phiKt*)
INST
INV
AC
GEN
/ PWR
PWR Y,/ ™
INV m
SEC m
t0AD Q -O I 0— 0 Kill
TRANSFER y
RELAY 0-0 O
CLOSE
TRIP
FIRE CONTROL
;if r i-
SINGLE-PHASE KSSTYTIAL BUS
Energized by main inverter
when switch is at NORMAL,
by spare inverter when
switch is at EMERGENCY.
AFTERBURNER SPEED SENSING SWITCH
AUTOPILOT
AUTOSYN INSTRUMENTS
CABIN TEMPERATURE CONTROL
ENGINE IGNITION
FUI1 QUANTITY SYSTEM
GLIDE SLOPE RECEIVER
SIDESLIP STABILITY AUGMENTER
VHF NAVIGATION RECEIVER
WINDSHIELD DE-ICE AND
DEFOG CONTROLLER
(AUXILIARY A C BUS WHIN
ALTERNATOR IS IN0PIRATIVU
Normally energized by single¬
phase essential bus
through phase converter.
Energized by three-phase inverter
when single-phase essential
bus is not energized.
FLIGHT COMPUTER
DIRECTIONAL INDICAT OR
Wm
/
_ To)
EXTERNAL POWER UNIT
AUXILIARY A-C BUS
Energized by single-phase
essential bus if alternator fails.
FIGHTER IDENTIFICATION SYSTEM
RADAR
WINDSHIELD DEFOG
three-phase inverter bus
Energized by either main or
spare three-phase inverter,
or by external power.
ALTERNATOR BUS
Energized by alternator or
external power unit.
Fm CONTROL
FUEL VINT DEICE HEATERS
MISSILE HEATERS
NADAR NEATER
RADAR
R A DOME ANTI-ICE FLUID HEATER
WINDSHIELD DE-ICE HEATER
H23(2)t
■■■■■
1-27
Section I
T.O. 1F-89H-1
Figure 1-22,
d-c generator voltage output can be increased or de¬
creased by turning the voltage regulator rheostats
toward INGR or DEC, The voltage regulators are
normally preset on the ground by qualified person¬
nel and should not be readjusted in flight unless an over¬
voltage condition exists which continually disconnects
the generator from the bus,
28-Volt D-C Generator Switches.
For each 28-volt d-c generator there is a guarded gen¬
erator switch {figure 1-22), located on the 28-volt d-c
control panel. The function of these switches is to
connect the corresponding generator to the 28-volt d-c
primary bus and to reset the field control relay after
an overvoltage condition has occurred- The switch
positions are ON, OFF, and RESET* The switch is
spring-loaded to OFF from the RESET position. Plac¬
ing the switch at ON connects the generator to the
primary bus; at OFF, it disconnects the generator
from the bus. The RESET position is used as follows:
If the voltage of a generator becomes excessive, an
overvoltage relay opens the generator field circuit and
causes generator voltage to drop to zero. As the volt¬
age drops, a reverse-current cutout relay disconnects
the generator from the primary bus. To return the
generator to service, the switch must be placed mo¬
mentarily at RESET. A circuit is then completed to the
generator field and generator voltage builds up to
normal. Then the switch can be placed at ON to
complete the circuit between the generator and the
28-volt d-c bus. If the overvoltage condition persists
(as indicated by the generator warning light again
1-28
T.O. 1F-89H-1
Section L
coming on)* voltage can be reduced to the correct
value by first placing the generator switch at OFF*
then turning the voltage regulator rheostat knob
toward DEC (counterclockwise). Next, the generator
switch must be placed momentarily at RESET, then
returned to OFF. With the switch at OFF, the voltage
regulator rheostat knob should be adjusted so that
the voltmeter reads 28 volts. Then the generator switch
can be placed at ON to put the generator back into
service.
28-Volt D-C Generator Warning Lights.
Each generator has a 28-volt d-c generator-off warning
light (figure 1-22) located on the 28-volt d-c control
panel. These lights are marked GEN OFF. The lights
come on to warn the pilot when the corresponding
generator is disconnected from the 28-volt primary
bus. The light will come on under the following con¬
ditions: before engines are started when the battery
switch is turned ON or an external source of d-c
power is applied to the airplane; when the engines
are operating but the generator switch is at OFF; or
if the generator has been automatically disconnected
because of an overvoltage condition.
28-Volt D-C Circuit Breakers.
Most of the 28-volt d-c circuits (except emergency
circuits) are protected by push-pull circuit breakers
{figure 1-25) on five circuit breaker panels: two on
the pilot’s left console, one on the pilot’s right con¬
sole, and one each on the left and right sides of the
radar observer’s cockpit. Electrical overload within
a circuit will cause the corresponding circuit breaker
to pop out and open the overloaded circuit. The cir¬
cuit may be closed again by pushing the circuit break¬
er IN, or the circuit can be opened manually by pull¬
ing the circuit breaker OUT.
28-Volt D-C Loadmeters,
Three loadmeters (figure 1-22), one for each gen¬
erator, are located on the 28-volt d-c indicator panel
on the pilot’s right console. The loadmeters indicate
the proportion of generator rated output being used.
28-Volt D-C Voltmeter and Voltmeter
Selector Switch*
A voltmeter and a voltmeter selector switch (figure
1-22), on the 28-volt d-c indicator panel on the pilot’s
right console, provide a means of determining gen¬
erator voltage output. The selector switch has LH
GEN, RH GEN NO. 1, RH GEN NO. 2, PRI BUS,
SEC BUS, and OFF positions. When the switch is
turned to one of the three generator positions, the
voltmeter indicates the output of the generator se¬
lected. When the switch is turned to PRI BUS or
SEC BUS, the voltmeter indicates the voltage being
supplied to the bus selected. When the switch is at OFF,
the circuits to the voltmeter are open and the volt¬
meter reads zero.
Pilot's right console
INVERTER
CONTROL PANEL
H-25B
Figure 1-23 .
Note
Whenever the engines are operating, the volt¬
meter will indicate a voltage from each 28-volt
d-c generator whether the generator switch
is at ON or at OFF, unless the generator field
circuit has been broken by action of the over¬
voltage relay or by generator failure. The
loadmeter, however, will indicate load only
when the generator switch is at ON and
power is being supplied to the 28-volt d-c
primary bus.
INVERTER SYSTEMS,
Alternating current is supplied by two 115-volt in¬
verter systems; a single-phase system and a three-phase
system. Each system has two inverters powered by
28-volt dc. In the single-phase inverter system, two
2500-va inverters (a main and spare) supply power
to the essential and secondary buses (see figure 1-21).
During normal operation, both single-phase inverters
operate; the main inverter supplies power to the essen¬
tial bus, and the spare inverter supplies power to the
secondary bus. All single-phase inverter powered equip¬
ment is protected by circuit breakers on a panel (figure
1-25) located on the bulkhead at the right aft side
of the pilot's seat. All inverters are powered by the 28-
volt d-c essential bus; however, the control circuit for the
T-29
Section E
T.O. 1F-S9H-1
K-SM
Figure 1 - 24 .
spare single-phase inverter receives its power from
the secondary 28-volt d-c bus, which is energized when
two or more 28-volt d-c generators are operating. If
the main single-phase inverter fails during normal
operation, a red warning light will come on to indi¬
cate that the essential bus is not energized. Then
emergency operation can be selected; the spare single¬
phase inverter, by means of a load transfer relay, will
power the essential bus and the secondary bus will not
be energized. If the spare inverter fails during normal
operation, the main inverter will continue to supply
power to the essential bus and the secondary bus will
not be energized. The essential bus, in addition to
carrying its normal load, also supplies power to the
auxiliary bus (normally powered by the alternator)
in case the alternator fails. In addition to equipment
operated directly from the essential bus, the gyrosyn
compass system and the flight computer are powered
through a phase converter by the essential bus. The
phase converter changes single-phase power to three-
phase power. If the essential bus is not energized, as
would occur if both single-phase inverters fail, a load-
transfer relay will automatically shift the load of the
gyrosyn compass system and the flight computer to the
three-phase inverter system. The three-phase inverters, a
main and a spare, are each rated at 500-va. Only one
three-phase inverter (main or spare) operates at a time.
Normally only the attitude indicator is powered by the
three-phase inverter system. Operation of either main
or spare three-phase inverter is manually selected. A
red warning light will come on to warn of either
three-phase inverter failure or an open attitude indicator
circuit breaker. All controls and indicators, except the
a-c voltmeter and voltmeter selector switch, for both
three-phase and single-phase inverter systems are on one
inverter control panel located on the pilot's right console.
The voltmeter and voltmeter selector switch, which serve
both inverter systems and the alternator system, are locat¬
ed in the radar observer’s cockpit on the alternator con¬
trol panel.
Single-Phase Inverter Switch,
A 28-volt d-c switch (figure 1-23) on the upper portion
of the inverter control panel has NORMAL, OFF, and
EMERGENCY positions to control single-phase inverter
operation. When the switch is at NORMAL, both single¬
phase inverters operate; the main inverter powers the
1-30
T.O. IF-89KM
Section 1
Pilot's right console
Pilot's left console
Section I
T.0* 1F-89H-1
essential bus and the spare inverter powers the secon¬
dary bus. When the switch is at EMERGENCY, the
spare inverter powers the essential bus, the secondary
bus is not energized, and the main inverter does not
operate. The EMERGENCY position is used only when
the main inverter fails. When the switch is at OEF,
both single-phase inverters are deenergized. Either in¬
verter can be operated individually by pulling the cir¬
cuit breaker for the other inverter.
Three-Phase Inverter Switch.
A 28-voit d-c switch (figure 1-23) on the upper portion
of the inverter control panel has MAIN, SPARE, and
OEF positions to control three-phase inverter operation.
When the switch is placed at MAIN or SPARE, a cir¬
cuit is completed from the 28-volt d-c bus to the
corresponding inverter* When the switch is at OFF,
both main and spare inverters are inoperative. A red
warning light burns if the selected inverter (main or
spare) is not operating, or if the inverter switch is at
OFF.
Single-Phase Inverter Warning Light.
A red warning light (figure 1-23) on the single-phase
portion of the inverter control panel indicates when
the essential bus of the single-phase inverter system
is not energized. The light is marked NO AC POWER
—ESSENTIAL BUS and operates on 28-volt dc. If the
light comes on while the single-phase inverter switch
is at NORMAL, the switch can be moved to EMER¬
GENCY so that the spare inverter will supply power
to the essential bus. As soon as the essential bus re¬
ceives power, the light will go out. The light will burn
when the switch is at OFF.
Three-Phase Inverter Warning Light.
A red warning light (figure 1-23) marked NO AC
POWER and located on the three-phase portion of
the inverter control panel comes on if the selected
(main or spare) three-phase inverter is inoperative,
if the inverter switch is at OFF, or if the attitude indica¬
tor circuit breaker is open. The light operates on 28-
volt dc.
Note
When the single-phase inverter switch is
moved from NORMAL to EMERGENCY, the
three-phase inverter light will flicker on mo¬
mentarily* This is a result of the three-phase
inverter momentarily picking up the gyrosyn
compass system and flight computer load while
the changeover is being made.
A-C Voltmeter and Selector Switch.
One voltmeter (figure 1-24) is provided for both the
inverter systems and the alternator system. The volt¬
meter and its selector switch (figure 1-24) are lo¬
cated on the radar observer’s alternator panel. (For a
complete discussion on the voltmeter and selector
switch, see paragraph entitled A-C Voltmeter and
Selector Switch included in subsequent discussion on
the a-c alternator system, this section,)
ALTERNATOR SYSTEM.
A variable frequency alternator, driven by the left en¬
gine, supplies three-phase 115/200-volt ac to two
buses: the alternator bus and the auxiliary a-c bus (see
figure 1-21). An exciter switch turns on the alternator
by energizing the alternator fields. An alternator cir¬
cuit breaker connects the alternator, through a relay,
to the two buses* Both switches must be placed mo¬
mentarily at CLOSE to obtain alternator output. Alter¬
nator failure will cause a bus-tie relay to connect the
auxiliary a-c bus to the essential single-phase inverter
bus*
Alternator External Power Switch.
The three-position external power switch (figure
1-24) on the radar observer's alternator control panel
controls the external power circuit breaker. The switch
is spring-loaded to NEUTRAL from the CLOSE and
TRIP positions* After a 115/200-volt 400-cycle a-c ex¬
ternal power source is connected to the external power
receptacle, the external power switch can be held
momentarily at CLOSE to close the circuit breaker
connecting the external power source to the distribu¬
tion bus* Holding the switch momentarily at TRIP
discontinues external a-c power to the distribution bus.
When the alternator circuit breaker switch is held to
CLOSE, it automatically trips the external power cir¬
cuit breaker.
Operation of more than one alternator switch
at a time will result in damage to the alter¬
nator control circuit.
Note
Before the external power switch can be
closed, 28-volt d-c external power must be
connected*
Alternator Exciter Switch.
Two three-position exciter switches (figure 1-24), one
on the pilot’s alternator control panel and one on the
radar observer's alternator control panel, control 28-
volt d-c circuits to the alternator exciter relay and
provide a means for either crewmember to turn the
alternator on and off. These switches are spring-loaded
to NEUTRAL from the CLOSE and TRIP positions.
When either switch is held momentarily at CLOSE,
a circuit is completed from the 28-volt d-c bus to the
exciter relay, which in turn closes and turns on the
1-32
T.O. TF-39H-1
Section I
alternator. When the switch is held momentarily to
TRIP, the circuit from the 28-volt d-c bus to the ex¬
citer relay is broken; the relay opens and cuts off
alternator output.
L CAUTION j!
Operation of more than one alternator switch
at a time will result in damage to the alter¬
nator control circuit.
Alternator Circuit Breaker Switch and
Indicator Light*
Two three-position circuit breaker switches (figure
1-24), one on the pilot’s alternator control panel and
one on the radar observer’s control panel, close or trip
the alternator circuit breaker. Each switch is spring-
loaded to NEUTRAL from the CLOSE and TRIP
positions. Holding the switch momentarily in the
CLOSE position closes the circuit breaker connecting
the alternator to the distribution bus and automat¬
ically trips the external power circuit breaker. Hold¬
ing the switch momentarily in the TRIP position
opens the circuit breaker, discontinuing alternator
output to the distribution bus. The red indicator light
(figure 1-24) to the right of the circuit breaker switch
in each cockpit comes on when the alternator circuit
breaker is in the tripped position.
| CAUTION
1
Operation of more than one alternator switch
at a time will result in damage to the alter¬
nator control circuit.
Alternator Voltage Rheostat.
A guarded voltage rheostat (figure 1-24) on the radar
observer’s alternator control panel can be used to ad¬
just the voltage output of the alternator.
A-C Voltmeter and Selector Switch*
A voltmeter and selector switch (figure 1-24), located
on the radar observer’s alternator control panel, are
used to check the voltage of all a-c power systems. The
rotary selector switch has OEF, EXT PWR, PWR
INV PRI, PWR INV SEC, AC GEN, INST INV,
and BUS positions. When the switch is at EXT PWR,
the voltmeter indicates external a-c power voltage be¬
fore the external power switch is closed. When the
switch is at PWR INV PRI or PWR INV SEC, the
voltmeter indicates the voltage of the essential or sec¬
ondary single-phase bus. When the switch is at AC
GEN, the voltmeter indicates the voltage output of the
alternator. When the switch is at INST INV, the volt¬
age indicated is that of the selected three-phase inverter
(main or spare). When the switch is at BUS, the
alternator bus voltage is indicated.
HYDRAULIC POWER SUPPLY SYSTEM.
The complete hydraulic power installation includes a
left system and a right system, both powered by engine-
driven pumps, with a supplemental electrically driven
hydraulic pump tied into the left system. No interflow
can occur between the left and right systems. The left
and right systems operate at 3000 psi, and the supple¬
mental hydraulic pump at 2500 psi. Each primary flight
control has two actuating cylinders: one powered by
the left system, and one powered by the right system.
If either the left or right system fails, the remaining
system provides adequate but limited flight controL
If both the left and right systems fail, the left hydrau¬
lic system supplemental pump provides further limited
flight control if the left hydraulic system has not failed
through loss of hydraulic fluid. One pressurized hy¬
draulic reservoir for the left system and one for the
right system are in the forward fuselage section. The
reservoirs are pressurized to prevent the fluid from
foaming at altitude and to maintain a positive pressure
on the inlet side of the engine-driven pumps. During
engine starts, a purge valve, one in each system, by¬
passes hydraulic fluid from the pump back to the
reservoir to reduce the load on the starter. After the
engine starts, the pump puts out more fluid than the
purge valve can bypass. The increase of pressure in the
valve overcomes a spring tension and forces a piston
over the return line to close the valve. System pressure
then builds up to 3000 psi. During cold weather, for
ground operation only, the hydraulic fluid in the left
and right systems is maintained automatically at operat¬
ing temperature. The weight of the airplane on the
landing gear energizes a circuit to a thermoswitch.
When the fluid temperature drops below a predeter¬
mined value, the thermoswitch actuates an electric
shutoff valve and the fluid is routed through a restric¬
tor which raises the temperature of the fluid until the
correct temperature is obtained. (See figures 1-26, 1-27,
1-30, 1-33, 1-36, and 1-37 for hydraulically operated
equipment. Refer to figure 1-45 for hydraulic fluid
specification.
LEFT HYDRAULIC SYSTEM*
Basic operating pressure for the left system comes from
an engine-driven piston-type hydraulic pump on the
left engine and an electrically driven supplemental
hydraulic pump. This system powers one actuating cyl¬
inder on each flight control surface, the landing gear,
main gear inboard doors, wheel brakes, wing flaps,
speed brakes, missile extension mechanisms, and the
nose wheel steering system. The left system includes a
1-33
Section 1
T*Q. 1F-89H-1
Hydraulic Power Supply Systems
PRESSURE (LEFT SYSTEM)
PRESSURE (RIGHT SYSTEM)
HANDPUMP PRE5SURE
SUPPLY
RETURN (LEFT SYSTEM)
V/i
RETURN {RIGHT SYSTEM)
COMPRESSOR AIR
ELECTRICAL ACTUATION
MECHANICAL ACTUATION
CHECK VALVE
&J
\<*t.
THERMAL SWITCH
LANDING GEAR STRUT SWITCH
LANDING GEAR HANDLE
RELIEF VALVE
{SPRING
PRESSURE SWITCH LOADED
OPEN)
PRESSURE
transmitter
7b brakes
FIGURE 1-37
7b rocket-missile pod system
To flight controls
FIGURE 1-27
To flaps
and speed brakes
FIGURE 1-30
H-27C
figure 1-2 6 .
1-34
T.O. 1F-89H-I
Section I
pressurized reservoir in the left side of the forward
fuselage section, a brake accumulator in the nose gear
wheel well, and a handpump and two selector valves
in the radar observer's cockpit. The handpump is or¬
dinarily used to operate the hydraulic engine hoist
system* In an emergency, the radar observer can re¬
charge the brake hydraulic accumulator by placing the
two selector valves at the proper placarded positions
and then actuating the pump handle.
Note
The engine hoist system includes two hydrau¬
lic cylinders m the aft fuselage section, one of
the two selector valves in the radar observer’s
cockpit, and needle control valves under the
aft lower wing fillet doors. The hoist system is
used by ground crew personnel when engine
service is required. The handpump will not
maintain sufficient hydraulic pressure for op¬
eration of the flight controls.
The left hydraulic system supplemental pump is started
in three different ways. It starts automatically either
in flight or on the ground whenever brake accumu¬
lator pressure drops below 1150 to 800 psi. A landing
gear lever switch also starts the pump automatically
when the landing gear lever is moved to the DOWN
position in flight, to supply an additional volume of
hydraulic flow to lower the gear* A strut switch cuts
out the landing gear lever switch to prevent pump
operation while the airplane's weight is on the gear.
Normally, in flight and during taxi operations, the
supplemental pump can be energized by depressing
the nose wheel steering button, and deenergized by
releasing the button. However, if the left hydraulic
system pressure switch is automatically actuated, be¬
cause of excessive use of the wheel brakes during
taxiing, the supplemental pump will be automatically
energized and continue to operate until the left hy¬
draulic system pressure reaches 2200 to 2350 psl,
regardless of the nose wheel steering button position.
The steering and brake systems have first priority on
supplemental pump flow and only the surplus flow
enters the main left hydraulic system. This provides
adequate flow on the ground for braking and steering
regardless of other hydraulic system functions, even
with the left engine inoperative. Since braking and
steering are not used in the air, all the flow enters
the left main system when the nose wheel steering
button is depressed or the gear lowered, providing the
brake accumulator is fully charged.
Nofe
In the event of a complete power failure, the
battery switch must be ON to operate the sup¬
plemental pump.
CAUTION ;i
• When a demand is made on the supplemental
pump by operation of any left hydraulic sys¬
tem control, the supplemental pump must not
be in operation for a period of more than
6 minutes, followed by a rest period of 15
minutes.
• When no demand is made on the supplemen¬
tal pump by operation of any left hydraulic
system control, the supplemental pump should
not be in operation for more than 30 minutes.
RIGHT HYDRAULIC SYSTEM.
Operating pressure for the right system is normally
supplied by an engine-driven piston-type hydraulic
pump on the right engine. This system powers one
actuating cylinder of each basic flight control surface.
The pressurized reservoir for the system is in the right
side of the forward fuselage section.
HYDRAULIC SYSTEM PRESSURE GAGES.
Both left and right systems and the brake accumulator
system have autosyn pressure gages (figure 1-11) on
the pilot's center pedestal* The gages operate on 115-
volt ac from the main or spare single-phase inverter.
A pressure gage (figure 1-9), showing the air pressure
in both left and right system reservoirs, is located
above the pilot's left console.
FLIGHT CONTROL SYSTEM.
Hydraulic actuating cylinders controlled by servo
valves operate the control surfaces of the airplane. The
servo valves are in turn controlled by push-pull rods
and cable linkages from the pilot's stick and rudder
pedals. The rudder has a single set of control cables,
and the elevator and ailerons have dual sets of control
cables. All control surfaces have two independent sets
of hydraulic actuators. One set receives hydraulic
power from the right hydraulic system and the other
from the left hydraulic system. Either system will give
adequate control for safe flight. Surfaces other than
the rudder operate on the 3000-psi system pressure.
The rudder actuating cylinders operate on 700-psi
pressure obtained through pressure reducers which re¬
duce the normal 3000-psi system pressure. Since the
flight control surfaces are fully powered, artificial
“fee! 1 * has been provided because no forces are trans¬
mitted to the stick and the rudder pedals, A bob weight
on the control force mechanism and a control force
bellows, utilizing ram air pressure, provide additional
"feel” for elevator operation. The irreversible surface
control hydraulic system opposes surface movement
when the airplane is not in use; however, the control
1-35
Section 1
T.O. 1F-89H-1
Control Hydraulic System
\
From right system
From left system
AILERON CYLINDERS
AND SERVO VALVES
PRESSURE REGULATING VALVE
RUDDER CYLINDER
AND SERVO VALVE
RUDDER CYLINDER
AND SERVO VALVE
ELEVATOR CYLINDER
AND SERVO VALVE
i
(
<
(
1-36
Figure 1 -27.
T.O. 1 F-89H-1
Section I
surfaces will eventually droop after the airplane is
parked without hydraulic pressure on the system. This
is normal and should cause no alarm, as the control
surfaces will return to their normal positions when
hydraulic power is applied.
CONTROL STICK.
The control stick (figure 1-28) is conventional with the
following 28-volt d-c switches on the grip: aileron
and elevator trim switch, pylon tanks and bombs re¬
lease button (inoperative), rocket-missile firing trigger,
radio mike button, autopilot emergency disconnect
switch, and nose wheel steering button which also ac¬
tuates the left system supplemental hydraulic pump.
RUDDER PEDALS.
The rudder pedals are the conventional suspended
type with toe-operated brake pedals. The pedals are
adjustable to the desired position.
Rudder PedaB Adjustment Crank.
A rudder pedal adjustment crank (figure 1-11) is
on the pilot's center pedestal panel. Rotation of the
crank moves both rudder pedals either forward or aft
to the desired position.
ELEVATOR FEEL SYSTEM.
A control force bellows in the elevator control mecha¬
nism lends "feel” for elevator movement in propor¬
tion to airspeed, A diaphragm in the bellows is at¬
tached so that a movement of the stick in either
direction moves the diaphragm against ram-air pres¬
sure, In flight, ram air from the right pitot head
creates the pressure on the diaphragm. This pressure in¬
creases with airspeed, increasing the resistance to con¬
trol stick movement. When the airplane is not moving,
there is no differential pressure in the bellows and no
bellows resistance to control stick movement; however,
elevator "feel” is provided by a spring within the
bellows. Additional feel on the control stick comes
from a bob weight attached to the stick mechanism.
When "G” forces are applied to the airplane, the bob-
weight tends to move the stick toward the position of
one "G” flight. The stick force increases as the “G”
force becomes greater.
FLIGHT CONTROL TRIM SYSTEM.
The control stick or pedal forces can be relieved by
use of the trim system. The ailerons and elevator are
trimmed by electric motors that mechanically change
the relationship between the "feel” mechanism and
the control system to reduce stick force to zero. The
trim system operates directly on the control force
AILERON AND ELEVATOR
TRIM SWITCH
{Inoperative
ROCKET-MISSILE
FIRING TRIGGER
MIKE BUTTON
NOSE WHEEL STEERS
BUTTON AND LEFT
SUPPLEMENTAL PUMP
ACTUATION
(SOME AIRPLANES)
F/gvre 1-28.
producers and no trim tabs are used on the control
surfaces. Aileron and elevator trim is accomplished by
moving the aileron and elevator trim switch on the
control stick grip. Limit switches are provided on rhe
elevator trim mechanism to prevent serious over trim
if the switch should stick. Aileron trim travel is 6
degrees each way from neutral. Elevator trim travel is
11 degrees up and 10 degrees down. The rudder is nor¬
mally trimmed automatically through rhe sideslip sta¬
bility augmenter. The rudder may also be trimmed
manually in emergencies by rotating the rudder trim
knob either left or right. The rudder can be manually
trimmed up to 5 degrees each way from neutral. Man¬
ual rudder trim should be used only when the sideslip
stability augmenter system is inoperative.
Aileron and Elevator Trim Switch.
The aileron and elevator trim switch (figure 1-28)
on the pilot’s control stick grip can be moved up or
down for elevator trim and left or right for aileron
1-37
Section I
T.0. 1F-89H-T
trim. This switch, operating on 2 8-volt dc, controls
electrical trim motors that reduce the stick force to
zero, within trim limits, at a chosen aileron or elevator
position.
manually to preclude the possibility of the
switch sticking in the actuated position and
causing a dangerous over trim condition in
case of malfunction of the limit switches.
The aileron and elevator trim switch is
spring-loaded to the NEUTRAL position;
however, it should be returned to NEUTRAL
HIOHT CONTROL TRIM SYSTEM
SmSUP STABILITY
AUGMENTER CONTROL PANU
(LEVATOR TRIM INDICATOR
Pilot's cockpit —
side
Figure 1 - 2 9*
Note
The ailerons and elevator cannot be trimmed
unless both hydraulic power and 28-volt d-c
electrical power are available*
Electrical Rudder Trim Knob.
A rudder trim knob (figure 1-29), located on the side¬
slip stability augmenter control panel on the pilot's left
console, provides a means of trimming the rudder man¬
ually. Hydraulic pressure, 115-volt single-phase ac, and
28-volt dc are required for effective use of the knob.
The knob is safetied in the NEUTRAL position and is
used only as an alternate means of trimming the rudder
in case of malfunction of the autotrim feature of the
sideslip stability augmenter. A rudder travel of 5 de¬
grees in each direction can be obtained by rotation of
the knob, which changes the position of the rudder
servo with respect to the normal pedal position. When
the rudder trim knob is rotated clockwise, the rudder
deflects to the right. When the rudder trim knob is
rotated counterclockwise, the rudder is deflected to the
left. To use this trim knob, the rudder trim switch is
moved to MANUAL TRIM position and the trim knob
is rotated, to the right or left as required, with suffi¬
cient force to break the light safety wire.
Rudder Trim Switch.
A rudder trim switch (figure T29), on the sideslip sta¬
bility augmenter control panel is for selecting either of
two methods of trimming the airplane directionally
through the sideslip stability augmenter system. This
switch operates on 28-volt dc and has positions marked
AUTO TRIM and MANUAL TRIM. When the switch
is at AUTO TRIM, the airplane is automatically kept
in directional trim. When the switch is moved to
MANUAL TRIM position, directional trim is accom¬
plished through a rudder trim potentiometer and the
rudder centering mechanism by turning the rudder
trim knob to the right or left as required.
Note
# The rudder cannot be trimmed unless hydrau¬
lic power, 28-volt d-c electrical power, and
115-volt essential bus power are available.
• At low indicated airspeeds, normally asso¬
ciated with takeoff, landing, and cruise at
extreme altitudes, a pressure switch (in the
landing gear warning system) overrides the
1-38
T.O. 1F-89H-1
Section I
Speed Smites and
Wing flaps Hydraulic System
PRESSURE
(LEFT SYSTEM)
OPEN
(SPEED BRAKES)
DOWN
(FLAPS)
CLOSE
(SPEED BRAKES)
UP
(FLAPS)
From left system
WING FLAP SERVO VALVE
WING FLAP HYDRAULIC MOTOR
H31C
Figure 7-30.
AUTO TRIM position of the rudder trim
switch. This action takes place automatically
when the airspeed drops below 165 knots
IAS; then the system returns to normal when
the airspeed builds up to 180 knots IAS, The
rudder trim switch itself does not move, as
the pressure switch is in sequence with it.
During the time that the autotrim feature
is not in operation, electrical manual trim
will be available through the sideslip sta¬
bility augmenter system just as though the
rudder trim switch were placed in MANUAL
TRIM position. This automatic switching in
and out of autotrim is to prevent undesirable
oscillations that might occur with the auto¬
trim feature operating at low indicated air¬
speeds*
1-39
Section I
T.O, 1F-S9H-1
Elevcttor Trim Position Indicator.
A mechanical elevator trim indicator (figure 1-29)
shows the proper trimmed position of the control
stick for takeoff. The indicator is located on the floor
at the inboard side of the pilot’s right console. The
indicator pointer is connected directly to the control
stick elevator torque tube and the dial is fixed to the
structure. The dial has a luminous circular spot
marked TAKE OFF, To trim the stick for takeoff, the
stick must be moved until the pointer is at TAKE
OFF* With the stick held in this position, the elevator
trim switch on the control stick grip must be ac¬
tuated until the stick force is reduced to zero. The
stick force should be reduced to zero within 10 sec¬
onds and the stick will remain at the indicated TAKE
OFF position,
SIDESLIP STABILITY AUGMENTER
SYSTEM*
The sideslip stability augmenter system controls rud¬
der motion to eliminate sideslip. This improves sta¬
bility, dampens undesirable oscillations (Dutch Roll)
common to most high speed airplanes, and permits
fully coordinated turns to be made without use of the
rudder pedals. If the airplane starts to sideslip, an
accelerometer of the mass-spring-damper type senses
the sideslip and transmits an electrical signal, propor¬
tional to the amount and direction of sideslip, to the
electronic control unit where it is amplified. The signal
is then transmitted simultaneously to the integrator and
the airspeed compensator. The signal to the airspeed
compensator is modified by an airspeed compensator
potentiometer and sent to the rate circuit where the
signal is again adjusted by a second airspeed com¬
pensator potentiometer. The signal is then combined
with a signal from the aileron potentiometer (propor¬
tional to aileron deflection) and sent through the sum¬
mer amplifier and a third airspeed compensator po¬
tentiometer for further modification. Signals from
the integrator and feedback potentiometer (propor¬
tional to rudder deflection) are combined with the
modified signal and transmitted to the power ampli¬
fier, The signal from the power amplifier controls an
electrohydraulic valve, that in turn controls the rate
and direction of hydraulic fluid flow to the rudder
power cylinders. The power cylinders then move the
rudder (without moving the rudder pedals) the
amount required to counteract the lateral acceleration.
The sideslip stability augmenter can be operated se¬
lectively either in automatic trim or in manual trim
at the pilot’s discretion. Automatic trim is recom¬
mended at all times and especially during the "on-tar-
get** stage of interception and the firing phase. In this
setting the system will produce the most stabilized
flight path at cruising speeds and above; however, the
system will provide satisfactorily stabilized flight and
is capable of continuous operation in the manual trim
setting. A sensitive air pressure switch (that opens at
165 knots and closes at 180 knots) is included in the
autotrim circuit, eliminating automatic trim at air¬
speeds below 165 knots. If the sideslip stability augmen¬
ter should fail completely in flight, the rudder may de¬
flect as much as 5 degrees either side of neutral (maxi¬
mum system authority). The rudder will return to neu¬
tral, however, within 60 seconds after the sideslip stabil¬
ity augmenter system is turned off. If a failure occurs in
the automatic trim portion of the electronic control unit
and power is still available to the system, or if the
E-ll autopilot (which automatically disconnects the
sideslip stability augmenter) is engaged, trim control
may be obtained through selection of the manual trim
system. The pilot may override the sideslip stability
augmenter at any time by use of the rudder pedals. This
system, powered by 115-volt ac, requires a warmup
period of approximately 30 seconds. During this
warmup period, the rudder may move as much as 5 de¬
grees either side of neutral. Therefore, if the system is
turned off during flight and then turned on again,
the rudder may deflect and the airplane will yaw
sharply. For this reason, the system should be turned
off in flighr only if there is complete system failure.
If the sideslip stability augmenter system
should fail, reduce airspeed below the range
in which large directional oscillations might
occur, thus avoiding undue stress on the air¬
plane’s structure.
Sideslip Stability Augmenter Power Switch.
A two-position PWR ON, PWR OFF switch (figure
1*29), located on the sideslip stability augmenter control
panel, controls the single-phase a-c power that operates
the sideslip stability augmenter system. The switch is
guarded in the PWR ON position and should be left in
that position at all times during flight unless the entire
sideslip stability augmenter system fails. If this occurs,
the switch should be placed at PWR OFF.
1-40
TO. 1F-89H-1
Section 1
Figure T-3I*
WING R AP SYSTEM,
The slotted wing flaps operate on hydraulic power
from the left hydraulic system {see figure 1-30). A wing
flap lever on the pilot's left console is connected by
cables to the wing flap servo valve mechanism which
controls the direction of fluid flow to a hydraulic mo¬
tor* Four jackscrew actuators, driven by the hydraulic
motor through a series of torque tubes, move the flaps
to the desired position. The flaps operate together.
Flap travel is 30 degrees down from the wing reference
plane* There is no emergency system for operating the
wing flaps; however, with the supplemental hydraulic
pump in operation, the flaps can be operated from this
pressure source if the left engine-driven hydraulic
pump fails.
WING FLAP LIVER AND POSITION INDICATOR.
The wing flap lever and position indicator (figure
1-31) are located on the pilot's left console. The lever
provides a means of moving the wing flaps to any
desired position and can be j>re-positioned at TAKE
OFF (flap 30 degrees down), DOWN (flap 50 de¬
grees fully down), and UP. As the wing flaps travel,
the indicator gives visual indication of the flap posi¬
tion at any time during travel. Although the wing flaps
can be pre-positioned only to the three detent positions,
they can be placed at intermediate positions by holding
the wing flap lever in the desired position until the in¬
dicator shows the flaps to be in that position. The lever
can then be released and the flaps will remain in posi¬
tion until the lever is moved again. Retraction of wing
flaps from the TAKE OFF to the UP position requires
approximately 10 seconds.
SPEED BRAKE SYSTEM,
The trailing section of each aileron splits through the
chord line to form two surfaces. The two surfaces,
hinged at the front, open to a V when used as a
speed brake. Each speed brake is operated by a hydrau¬
lic cylinder powered by the left hydraulic system. Flow
to the cylinders is regulated by the speed brake lever in
the pilot's cockpit through cables and servo valves.
Speed brakes may blow open if the airplane is parked
in a tailwind when external speed brake locks have not
been installed. There is no emergency system for operat¬
ing the speed brakes; however, with the supplemental
hydraulic pump in operation, the speed brakes can be
operated from this pressure source if the left engine-
driven hydraulic pump fails.
1-41
Section I
T.O. 1F-89H-U
SPEED BRAKE LEVER.
The speed brake lever (figure 1-32), located on the
pilot s left console, has OPEN and CLOSED positions
and controls the position of the speed brakes. When the
speed brake lever is moved, the speed brakes open to¬
gether proportionally to lever movement. The lever
can be stopped at any point between OPEN and
CLOSED to give intermediate positioning of the speed
brakes. At indicated airspeeds up to approximately 260
knots, the speed brakes can be fully opened (120 de¬
grees included angle). At indicated airspeeds above 260
knots, the lever can be pre-posirioned at any setting so
long as the lever is moved toward a more fully open
position, but the angle to which the speed brakes will
open will be decreased in proportion to the increased
airspeed. The speed brakes cannot be pre-positioned
toward the CLOSED position. The speed brake lever
must be pushed forward manually as the speed brakes
close. If airspeed is above 260 knots, the airload on the
speed brakes applies back pressure on the actuating
cylinders in excess of the hydraulic system pressure and
prevents full opening of the speed brakes. As airspeed
is reduced, speed brakes will open to the position pre¬
set by the lever,
LANDING GEAR SYSTEM.
The airplane has a tricycle landing gear which oper¬
ates on power from the left hydraulic system and is
controlled in normal operation by the landing gear
lever in the pilot’s cockpit. The main gear retracts
inboard into the wing and the nose gear retracts ver¬
tically into the fuselage. A selector valve, a sequence
valve, and actuating cylinders extend and retract the
landing gear and the main landing gear inboard
doors. The selector valve, attached by mechanical link¬
age to the pilot’s landing gear lever, directs the flow
of hydraulic fluid in the actuating cylinders to raise
and lower the landing gear and operate the main
landing gear inboard doors. The sequence valve re¬
verses the action of the hydraulic pressure in the ac¬
tuating cylinders of the inboard doors to synchronize
the opening and closing of the inboard door with
the retraction and extension of the main landing gear.
If the pressure in the left hydraulic system drops be¬
low 1450 psi, a priority valve closes to give the flight
control system priority over the landing gear system
by shutting off the flow of fluid through the landing
gear selector valve. Independent air bungee systems
aid normal and emergency extension of the landing
gear. Landing gear extension or retraction normally
takes 6 seconds; however, when the engine rpm is
below 80%, additional time may be required. The
pilot can reverse the normal landing gear cycle
at any time with a reverse movement of the
landing gear lever. Hydraulic pressure is automaticallv
relieved when all landing gear components are up
and locked; a hydraulic shutoff valve, spring-loaded to
open, operating on 28-volt dc, and controlled by micro-
switches, closes and shuts off the hydraulic pressure
to the selector valve. Pressure is reapplied if any up-
locks accidentally open during flight. On airplanes
SPUD BRAKC tern
SPEED BRAKE LEVE
SIGHT COMPUTER
CAGING BUTTON
MANUAL RANGING KNOB
FRICTION CONTROL
fSOME AIRPLANES)
(
1-42
Figure J-32.
T.O. 1F-89H-1
Section 3
Landing Gear Hydraulic System
To left system
Ew supplemental pump
MAJN GEAR ACTUATING CYLINDER
MAIN GEAR ACTUATING CYLINDER
H34C
figure 1 - 33 .
modified in accordance with T.O. 1F-89-639, the land¬
ing gear system solenoid shutoff valve has been re¬
moved, allowing full system pressure to be applied to
the main gear, nose gear, and main gear inboard doors
at all times. This ensures that the inboard main landing
gear doors are completely up and locked before the
main landing gear indicators indicate an UP and
locked position. When retracted, the landing gear is
completely enclosed by doors. The nose gear doors are
operated mechanically by the nose gear truss. Each
main gear outboard door moves with the stmt. Each
main gear inboard door is operated hydraulically by
two actuating cylinders and the sequence valve; the
door closes and locks after the main gear is extended.
If the landing gear lever is moved from one position to
the other before completion of extension or retraction,
a transfer piston on the sequence valve moves the
sequence valve to keep the main landing gear inboard
door open until the gear completes its movement in the
changed direction.
1-43
Section 1
T,Q. 1F-89H-1
Note
Ail airplanes have a controlled failure nose
landing gear drag brace pin and a reinforced
pilot's cockpit floor, A wheels down emer¬
gency landing is permitted regardless of ter¬
rain, which lessens the danger of personal
injury to the pilot if the airplane overruns the
runway during a landing or an aborted take¬
off.
Emergency Landing Gear System.
The emergency landing gear system allows gear exten¬
sion without hydraulic pressure. The emergency re¬
lease for the landing gear is a cable linkage from the
emergency release handle in the pilot’s cockpit to the
landing gear and doot uplocks. To prevent a fluid
lock in the gear cylinders, the normal landing gear
lever must be placed at the DOWN position before
the emergency release handle is pulled. Pulling the
handle will release the nose gear door locks, the nose
gear up lock, the main gear uplocks, and the main gear
inboard door locks. The landing gear will extend of
its own weight and be forced into the down and
locked position by the bungee system.
Landing Gear Ground Locks.
Ground safety locks (figure 1-34) are provided for
the main landing gear. The main gear locks are in¬
stalled between the hinge end of the lower side brace
and the point where the actuating cylinder attaches to
the strut. The nose gear ground lock is a clip which
slips over the downlock cylinder and is pinned in place.
All ground locks have red streamers attached.
LANDING GEAR LEVER.
The landing gear lever (figure 1-35), located on the
pilot's left vertical console, is mechanically linked to
the landing gear selector valve. The lever knob contains
GROUND SAFETY LOCKS
1-44
Figure 1 - 34 ,
H-35B
T.O. 1F-89H-I
Section J
H-36B
Figure 1-35,
a red light that indicates an open gear door or an
unsafe gear position for landing. When the lever is
moved to the DOWN position, the nose gear door locks
are opened mechanically and slightly in advance of the
nose gear up lock. The nose gear uplock switch for the
electrically operated hydraulic shutoff valve opens the
circuit to the shutoff valve. The valve opens and allows
system pressure through to the priority valve and the
selector valve. On airplanes modified in accordance
with T.O, IF-89-639, the hydraulic shutoff valve has
been removed, allowing full system pressure to be
applied to the main gear, nose gear, and main gear
inboard doors at all times. Hydraulic power is supplied
simultaneously to the actuating cylinders of the nose
gear, main gear, and inboard doors. As the main land¬
ing gear inboard door cylinders are compressed, the
door locks release and the door opens, releasing the
outboard door locks. Final movement of the inboard
doors releases main gear uplocks through a cable sys¬
tem. Final movement of the main gear actuates the
sequence valves and reverses the flow of fluid to the
inboard door actuating cylinders, causing the inboard
door to close and lock. As the nose gear extends, the
nose gear doors are moved to the open position. When
the landing gear lever is moved to the UP position, the
nose gear down lock releases. The inboard door cylin¬
ders compress, releasing inboard door locks and open¬
ing the doors. As the doors reach the open position, the
main landing gear actuating cylinders release the main
gear downlocks and all three landing gear actuating
cylinders simultaneously retract the nose and main
gears. As the nose gear retracts, the nose gear truss en¬
gages the nose gear door operating mechanism and the
doors close and lock. The main gear outboard door is
closed by the action of the main gear shock strut. As the
main gear enters its uplocks, the sequence valve is actu¬
ated and reverses the flow of fluid to the inboard door
actuating cylinders, closing and locking the doors. When
all the doors lock, the microswitches for the hydraulic
shutoff valve are actuated. The circuit is energized and
T-45
Section I
T.O, IF-89H-I
the shutoff valve closes, relieving the pressure in the
system. When the weight of the airplane is on the gear,
a solenoid plunger safety lock in the landing gear lever
quadrant automatically prevents accidental movement
of the gear lever to the UP position.
LANDING GEAR EMERGENCY RELEASE HANDLE.
The landing gear emergency release handle (figure
1*35), located on the pilot’s left vertical console, is
provided to lower the landing gear when the normal
system fails. Before the emergency release handle is
pulled, the landing gear lever must be placed at
DOWN. When the emergency release handle is pulled
to its full limit of travel (approximately 1 4 inches),
the locks for the main gear inboard doors and the nose
gear doors, and the uplocks for the main and nose
gear are opened mechanically by the cable system.
The landing gear extends of its own weight and is
forced into the down and locked position by the air
bungee systems. As the main gear extends, it pushes
the inboard doors open, and the doors remain open
until hydraulic pressure is again applied to the system.
The emergency release handle requires a hard pull of
approximately SO pounds to release the locks. The
pilot can feel each set of main gear locks release; first
the right gear, then the left. The nose gear will not
be felt as it is unlocked by the downward movement
of the landing gear lever. After the gear is down the
emergency release handle must be guided back to the
stowed position to prevent whipping. Since use of the
emergency system does not affect the operation of the
normal system, no readjustments are necessary after
the landing gear has been lowered by the emergency
system; as soon as hydraulic pressure is available the
gear may be operated by the normal method if the
malfunction was temporary.
LANDING GEAR EMERGENCY OVERRIDE LEVER,
If it is necessary to retract the landing gear with the
airplane on the ground, or if the solenoid plunger
safety lock fails, an emergency override lever (figure
1*35) releases the lock. When the airplane is on the
ground, the safety lock holds the landing gear lever
in the DOWN position to prevent accidental retrac¬
tion of the landing gear. The lock is automatically
retracted when the weight of the airplane is off the
wheels. The gear lever can be released in an emergency
by bolding the override lever down while moving the
gear lever up.
LANDING GEAR POSITION INDICATORS.
A landing gear position indicator (figures 1-8 and 1*35)
on the pilot’s instrument panel shows the position of
each gear. When a gear is down and locked, a wheel
will show in a small window corresponding to that gear.
When a gear is up and locked, IIP will appear in a
window. If a gear is unlocked or in an unsafe condi¬
tion or if the 28-volt d-c power is off, red and cream
diagonal stripes will show. The indicator tabs give the
position of the gears only; they are not controlled by
the gear doors. A red light in the landing gear lever
knob, operating on 28-volt dc, will come on and stay
on for any unsafe condition of the landing gear or
landing gear doors. The light will also come on any
time the warning horn is sounding, and will stay on
to indicate the gear is not down and locked even
though the warning horn is shut off by the reset
switch. If the light is indicating that the gear is not
down at low airspeed and low altitude, it will go off
when either airspeed or altitude is increased. The land¬
ing gear lever warning light will remain on until the
inboard main gear doors are retracted, even if the gear
is safe. For this reason, the gear position indicators and
a visual check for safe main gear should be relied upon
following emergency gear drop. On airplanes modified
in accordance with T.O. 1F-89-639, microswitches have
been installed between the aft main gear door locks in
the left and right gear wells to ensure that both in¬
board main landing gear doors are locked before the
main landing gear indicators indicate an UP and
locked position. On airplanes so modified an unsafe
condition will be shown on the landing gear indicators
if both the main gear and the inboard doors are not
up and locked.
LANDING GEAR WARNING HORN
AND RE5ET BUTTON.
The landing gear warning horn will give a steady
signal and the landing gear warning light will come
on if one or more of the landing gears are not com¬
pletely down and locked when the airspeed drops to
165 knots* plus or minus 10 knots. An altitude-sensing
switch prevents a warning signal at altitudes above
10,000 to 13,000 feet* depending on atmospheric con¬
ditions. A landing gear warning horn reset button
(figure 1-17) on the pilot's aft miscellaneous control
panel can be pressed to shut off the horn. The warning
system will be recycled if either the altitude or the
airspeed is increased above the warning range or if the
landing gear is extended. If the pilot does not use the
reset button, the horn will stop blowing automatically
when the airspeed reaches approximately 175 knots.
On airplanes modified in accordance with T.O. 1F-89-
627, the landing gear warning horn has been removed
and replaced with an audible warning signal unit.
If the landing gear has not extended and locked prop¬
erly on airplanes so modified, a warning signal will be
audible over the pilot’s headset. Operation and control
of the audible warning signal unit is the same as for
the landing gear warning horn which it replaces.
Note
A quick check of the indicator light in the
landing gear lever knob can be made when
the gear is down and locked. Pressing the
warning horn reset button will cause the in¬
dicator light to come on.
1-46
T,0, 1F-89H-1
Section I
HOSE WHEEL STEERING SYSTEM.
The dual nose wheel is equipped with a steering system
controlled by rudder pedal action, (See figure 1-36.)
The purpose of the system is to provide directional
control during taxiing and takeoff only. Hydraulic
pressure for the system is controlled by a solenoid shut¬
off valve operated by a button on the control stick grip.
When the shutoff valve is open, a servo valve, me¬
chanically controlled by the rudder pedals, directs pres¬
sure, according to the direction of rudder pedal dis¬
placement, to a vane-type actuator which turns the
nose wheel strut. A mechanical followup device returns
the servo valve to neutral when the nose wheel reaches
the displacement dictated by rudder pedal deflection.
The first 50 percent of rudder pedal displacement
causes the nose wheel to rotate only 6 degrees from
center. The remaining 50 percent of rudder pedal travel
rotates the nose wheel through the remaining 40 de¬
grees of angular displacement. When the nose wheel
steering system is not being used (shutoff valve dosed),
a bypass valve is open to permit free flow of hydraulic
fluid between both sides of the vane-type actuator, thus
allowing the nose wheel to swivel. Both steering and
swivel range of the nose wheel is 46 degrees each side
of the centered position, A limit switch on the nose
gear scissors closes the shutoff valve and opens the by¬
pass valve when the weight of the airplane is taken off
the nose gear strut, allowing it to extend. This allows
the nose gear to swivel so that the centering cam will
center the nose wheel for landing gear retraction and
extension. Nose wheel steering may be selected at any
time during taxiing and takeoff (assuming that the
weight of the airplane is on the nose wheel regardless
of the relative positions of the nose wheel and rudder
pedals. If the nose wheel position does not correspond
with the position of the rudder pedals when nose wheel
steering is selected, the nose wheel will turn to corre¬
spond to the rudder pedal position. The system operates
on pressure from the left hydraulic power supply sys¬
tem, Electrical components are powered by the 2 8-volt
d-c bus.
NOSE WHEEL STEERING ANED SUPPLEMENTAL
HYDRAULIC PUMP BUTTON.
A spring-loaded nose wheel steering button (figure
1-28) on the control stick grip controls the 28-volt d-c
shutoff valve and the actuator bypass valve in the hy¬
draulic steering system and the left hydraulic system
supplemental pump. When the button is pressed, the
shutoff valve opens, the bypass valve closes, the supple¬
mental pump starts, and hydraulic pressure is supplied
to the system. Subsequent movement of the rudder
pedals will then turn the nose wheel in the direction
and to the degree desired. The button must be held
Nose Wheel Steerinq
Hydraulic System
SYSTEM SHOWN IN OPERATING CONDITION
(Steering button depressed and nose wheel strut compressed)
Left system return
From left system
SHUTOFF VALVE
(Normally closed)
SERVO VALVE
PRESSURE (LEFT SYSTEM
RETURN (LEFT SYSTEM)
RIGHT TURN PRESSURE
(LEFT TURN RETURN
LEFT TURN PRESSURE
(RIGHT TURN RETURN
ELECTRICAL ACTUATION
MECHANICAL ACTUATION
(Normally open )
STEERING ACTUATOR
( Vane type)
nose gear strut SWITCH (Closed when
-9 strut is compressed)
BTgVSS VALVE Vs r
^ To supplemental
- hydraulic pump
NOSE WHEEL STEERING SWITCH
(Closed when
nose wheel
steering button
is depressed)
From 2H-volt d-c bus Im
Figure J-36.
Section I
T.O. IF-S9H-1
depressed during nose wheel steering operation. When
the button is released, the shutoff valve closes, the
bypass valve opens, and the nose wheel swivels freely.
A limit switch on the nose gear scissors overrides the
steering button and prevents the steering system from
operating when the weight of the airplane is not on
the nose gear. However, pressing the button in flight
will still start the supplemental pump to augment left
hydraulic system pressure.
BRAKE SYSTEM.
The main gear wheel brakes operate hydraulically
using pressure from the left hydraulic system and brake
accumulator which is pressurized by the left hydraulic
system. The power brake valves, operated by depress¬
ing the brake pedals, individually meter fluid to the
wheel brakes. If the left hydraulic system fails, brakes
can be operated a limited number of times by the pres¬
sure in the brake accumulator. In an emergency when
the accumulator pressure is gone, an emergency air¬
brake is available. (See figure 1-370 A normally open
pressure switch in the brake accumulator closes when
pressure drops to between 1100 to 800 psi (or below).
The switch starts the supplemental pump to replenish
braking pressure. On airplanes modified in accordance
with T.O- 1F-89H-522, an antiskid braking device is
incorporated in the brake system. This device is de¬
signed to allow maximum braking efficiency during
normal and adverse weather conditions without skid¬
ding the main wheels. For detailed discussion of wheel
brake operation see Section YII.
Note
Enough hydraulic brake pressure for parking
or towing can be obtained by operating the
hydraulic handpump in the radar observer’s
cockpit (figure 4-8).
Emergency Airbrake System.
If the normal hydraulic brake pressure is lost, a 1500*
psi storage bottle in the nose wheel well contains
enough air for at least three complete brake applica¬
tions. Turning the emergency airbrake handle to ON
and then pressing a brake pedal operates an airbrake
valve and meters air through a shuttle valve to the
wheel brake. The shuttle valve closes the hydraulic line
to prevent air from going into the hydraulic system.
Note
• Artificial 'Teel” is lighter for the emergency
airbrake system than for the normal hydraulic
brake system; therefore, when using the emer¬
gency system, anticipate greater braking ac¬
tion for a given pedal pressure.
• If both emergency airbrake and brake accumu¬
lator pressures are applied to the system simul¬
taneously, more pedal pressure will be re¬
quired for the same amount of braking because
the artificial “feel” for both systems must be
overcome at the same time,
BRAKE PEDALS.
The brake pedals are the conventional, toe-operated
type. Each pedal controls a hydraulic power brake valve
and an air power brake valve. When the pedals are
pressed, all four valves open and either air or hydraulic
pressure, or both, supply the braking action to the
wheels. "Feel” will be absent unless pressure is avail¬
able to one of the power brake valves. Application of
both air and hydraulic pressure increases the pedal
pressure required to obtain the same braking result.
PARKING BRAKE LEVER.
The parking brake lever (figure 1-11) is located on the
pilot's center pedestal. Pulling up on the parking brake
lever while depressing the brake pedals sets the parking
brakes. The parking brakes are released by manually
releasing the parking brake lever slowly while de¬
pressing the brake pedals.
EMERGENCY AIRBRAKE VALVE HANDLE.
The emergency airbrake valve handle (figure 1-9) is on
the pilot’s left console. Turning the handle to ON, and
then depressing the brake pedals, meters air to the
wheel brakes.
INSTRUMENTS.
Note
This paragraph covers only those instruments
which cannot be considered part of a complete
system.
The free air temperature gage and the turn and slip in¬
dicator operate on 28-volt dc. All the gyro-type instru¬
ments except the turn and slip indicator operate on 115-
volt three-phase ac. The standby magnetic compass, a
self-contained unit of conventional type, is suspended
from the windshield structure above and to the right
of the pilot's instrument panel. This magnetic compass
serves as a standby directional indicator in case the
directional indicator (slaved) or the 28-volt d-c power
fails.
INSTRUMENT PANEL VIBRATORS.
An instrument panel vibrator on the pilot's and radar
observer's instrument panels prevents the instruments
from sticking. Each unit, a miniature 28-volt d-c motor
driving an eccentric weight, operates continuously
when die 28-volt d-c power is on and the circuit breaker
is closed.
1-48
T.O. 1F-89H-1
Section I
Brake Hydraulic
and Air Systems
■■■
PRESSURE (LEFT SYSTEM)
SXKE
HANDPUMP PRESSURE
KO
RETURN (LEFT SYSTEM) 1
■ j
COMPRESSED AIR
AIR EXHAUST 'jj
EH
CHECK VALVE £<
ELECTRICAL ACTUATION f
n
—
MECHANICAL ACTUATION
GO=DQ
SHUTTLE VALVE
mss
RELIEF VALVE
E-ODD
HYDRAULIC PRESSURE SWITCH,
a!
From left system
supplemental pump
AIR FILLER VALVE
tram
handpump
BRAKE ACCUmUU-DR
AIR FILLER VALVE
\\ AIR POWER BRAKE VALVE
j „ ,
H-asc
Figure f-37*
1-49
Section I
T.O. 1F-89H-T
H-39C
Figure 1-38.
MACHMiTER.
A type A-l machmeter (figure 1-8), with the dial grad¬
uated in tenths and hundreths Mach, is on the pilot’s
instrument panel. The pointer indicates, regardless of
altitude and ambient temperature, the Mach number
at which the airplane is being flown. Numbers on the
dial run in tenths from 0.3 to 1.0 (below Mach 0.3 the
graduations are omitted because in this low-speed range
the airspeed indicator provides a more useful reference).
AIRSPEED INDICATORS.
The pilot’s airspeed indicator (figure 1-8) is calibrated
in knots and has mo pointers: a fluorescent pointer that
indicates airspeed and a red pointer with alternate bands
of fluorescent white that shows the airspeed that corre¬
sponds to a preset Mach number for the existing alti¬
tude. Clockwise movement of the red pointer is limited
by a stop which is preset at the limiting structural air¬
speed of the airplane. When the two pointers meet, the
airplane is flying at the maximum allowable airspeed
or the maximum allowable Mach number, whichever is
less. The upper half of the indicator dial contains a
window exposing a drum, graduated in 2-knot divisions
and geared to the main indicator pointer so that the
indicated airspeed can be read accurately to within 1
knot. The radar observer’s airspeed indicator (figure
4-6) is calibrated in knots and shows true airspeed. In
the true airspeed indicator a temperature-sensing bulb
and an altitude diaphragm automatically compensate
for temperature and altitude variations that affect the
airspeed reading,
ALTIMETER.
The pilot’s altimeter displays barometric pressure indi¬
cations in feet of altitude calibrations and is located
on the pilot's instrument panel (figure 1-8). The
altimeter has two hands, a notched disk with a pointer
extension, two setting marks, a warning indicator and
a barometric scale with an adjustment knob. The
longer of the concentrically arranged hands indicates
feet in units of 100, the shorter hand indicates feet in
units of 1000, and the notched disk with a pointer
extension indicates feet in units of 10,000. A warning
indicator consisting of a striped (cross-hatched) sector
painted on a dial above numeral five appears through
the notched disk at altitudes below 16,000 feet. An
outer setting mark indicating feet in units of 100 and
an inner setting mark indicating feet in units of 1000
operate in conjunction with the barometric scale and
are used when the pressure to be read is outside the
limits of the barometric pressure scale. The adjustment
knob is used to adjust the hands, setting marks, and
barometric scale simultaneously to correct for atmos¬
pheric pressure changes caused by changing climatic
conditions.
ACCELEROMETER.
A type B-6 accelerometer (figure 1-S) on the pilot’s
instrument panel indicates both positive and negative
accelerations. The accelerometer has three pointers.
The main pointer indicates existing accelerations. The
two auxiliary pointers stop at the highest acceleration
that has been reached; one indicates maximum positive
acceleration, and one indicates maximum negative
acceleration. A knob on the front of the instrument
case is used to reset the auxiliary pointers to zero. Until
they are reset, the auxiliary pointers will show the
maximum plus and minus movements of the main
pointer.
ATTITUDE INDICATOR.
A type B-1A attitude indicator (figure 1-38) on the
pilot’s instrument panel indicates the airplane’s atti¬
tude with respect to an artificial horizon. The instru¬
ment obtains d-c power from the primary bus and a-c
1-50
T.O. 1F-89H-1
Section I
power from the instrument inverter. The B-1A indi¬
cator is noncaging and incorporates a zero-pitch trim
knob that positions both the sphere and the horizon
bar to the zero position. The pitch trim knob has a
triangular mark for zero-pitch trim, three dots corre¬
sponding to a one-half inch deflection in the down¬
ward direction, and six dots corresponding to a 1-inch
deflection in the upward direction. The indicator has
a followup rate of 1 BO degrees per second in the pitch
and bank axis. The indicator has a fast initial erection
period, approximately 2 minutes ±30 seconds; but
if the indicator tumbles in flight, erection may take
15 minutes. Included in the indicator is an electrically
driven power warning flag that disappears from view
when the indicator is up to full speed and the system
is ready for operation. The flag will appear in case of
a complete ac or dc power failure. However, a slight
reduction in ac or dc power or failure of certain
electrical components within the system will not cause
the flag to appear, even though the system is not
functioning properly. The instrument operates through
360 degrees of roll and through 164 degrees of pitch.
The instrument is compensated for turn errors; how¬
ever, the lower sensitivity limit of the turn-error com¬
pensating mechanism is 40 degrees oer minute* Any
turn made below 40 degrees per minute will result in
turn errors common to other instruments. Turns made
above 40 degrees per minute will be compensated for
turn errors. In level flight, the maximum error in the
indication of the airplane’s attitude is less than one-half
degree.
© It is possible that a malfunction of the attitude
indicator might be determined only by check¬
ing it with the directional indicator (slaved)
and the turn and slip indicator*
@ A slight amount of pitch error in the indica¬
tion of the type B-1A attitude indicator will
result from accelerations or decelerations. It
will appear as a slight climb indication after
a forward acceleration and as a slight dive in¬
dication after deceleration when the airplane
is flying straight and level. This error will be
most noticeable at the time the airplane breaks
ground during the takeoff run. At this time a
climb indication error of approximately one
ENGINE FIRE
SELECTOR SWITCHES
H-40B
me
EXTINGUISHING
SYSTEM
AGENT DISCHARGE SWITCH
Figure 1-29*
1-51
Section ]
T.O, JF-B9H-1
and a half bar widths will normally be ob¬
served; however, the exact amount of error
will depend upon the acceleration and elapsed
time of each individual takeoff. The erection
system will automatically remove the error
after the acceleration ceases.
• If the power supply to the attitude indicator
is interrupted, the instrument will be un¬
reliable for 1 minute.
EMERGENCY EQUIPMENT.
FIRE EXTINGUISHING SYSTEM.
The fire extinguishing system has overheat detectors
and fire detectors in each engine nacelle, and a single
bromochloromethane extinguisher bottle in the nose
wheel well with a discharge line to each engine. Two
electrically fired, cartridge-operated, release valves and
a pressure gage are assembled on the bottle. When
either engine fire selector switch is placed in the up
position, all fuel valves necessary to isolate the affected
engine from its fuel supply close and the electrical
circuit for the fire extinguishing system is armed.
When the agent discharge switch is moved to ON,
current flows to the selected discharge valve on the
bottle and fires the cartridge which pierces a frangible
disk. The bottle discharges its entire contents into the
manifolding of the selected engine; the agent vaporizes
and so dilutes the oxygen content of the air in the
engine bay that it will no longer support combustion.
If both fire selector switches are actuated before the
agent discharge switch is actuated, the charge will be
distributed to both engines but it will be insufficient
to put out the fire in either engine. Both the fire
extinguishing system and its controls operate on power
from the 28-volt d-c bus. Overheat lights, fire warning
lights, and controls for the extinguisher are located on
a fire control panel on the pilot's right vertical console.
@ Repeated or prolonged exposure to high con¬
centrations of bromochloromethane (CB) or
decomposition products should be avoided.
CB is a narcotic agent of moderate intensity
but of prolonged duration. It is considered to
be less toxic than carbon tetrachloride, methyl
bromide, or the usual products of combustion.
In other words, it is safer to use than previous
fire extinguishing agents. However, normal
precautions should be taken including the use
of oxygen when available.
® This is a "one-shot'" fire extinguisher system.
The bottle must be replaced after use.
Fire and Overheat Warning Lights and Test Switch.
Two red fire warning lights (figure 1-39), one for each
engine, are located on the fire control panel and will
come on when a rapid temperature rise occurs in the
engine area. Two amber overheat warning lights
(figure 1-39), one for each engine, are on the fire
control panel and will come on when the tempera¬
ture in the engine bay rises above I78°C (350°F).
A single detector test switch (figure 1-39),spring-loaded
to an unmarked off position and with marked posi¬
tions, L & R FIRE CKT 1 and L OVERHEAT, and L &
R FIRE CKT 2 and R OVERHEAT, is for checking the
two fire and two overheat warning circuits. When this
switch is held at L & R FIRE CKT 1 and L OVERHEAT,
both fire warning lights should come on indicating that
fire warning circuit No. 1 is operative on both engines,
and the left overheat warning Light should come on
indicating that the overheat detectors in the left engine
bay are operative. When the switch is held at L &: R
FIRE CKT 2 and R OVERHEAT, both left and right
fire warning lights again should come on indicating
that fire warning circuit No. 2 is operative on both
engines and the right engine overheat warning light
should come on indicating that the overheat detectors
in the right engine bay are operative. When the cir¬
cuits are being tested, the ovet heat lights should come
on immediately; the fire warning lights, after a 2- to
10-second delay. The warning lights, test switch, and
detector circuits operate on 28-volt dc.
Engine Fire Selector Switches.
Two guarded fire selector switches (figure 1-39), one
for each engine, are mounted on the fire control panel.
These switches are used to turn off fuel shutoff valves
to the engine and to arm the fire extinguishing agent
discharge switch. When the guards over the switches
are down, the 28-volt d-c circuits to the agent discharge
switch and the fuel shutoff valves are broken. The guard
must be raised and the switch moved up to close fuel
valves for the affected engine and to complete the circuit
to the agent discharge switch.
Agent Discharge Switch.
A spring-loaded agent discharge switch (figure 1-39)
located on the fire control panel operates the fire
extinguisher. When the switch is held momentarily to
the ON position, the circuit is dosed and current flows
to the selected discharge valve on the fire extinguisher
bottle. There, a cartridge is fired to pierce a sealing disk,
and the full charge of extinguishing agent is directed
to the area surrounding the selected engine.
The agent discharge switch is ineffective
(unarmed) unless one of the engine fire selec¬
tor switches has been actuated.
T-52
T.O. 1F-89H-1
Section l
O RADAR OBSERVER S
CANOPY SWITCH AND
LOCKING LEVER
EXTERNAL
CANOPY HANDGRIP
RADAR OBSERVER S
CANOPY SLOW F3RE
JETTISON "T” HANDLE
RESCUE
INTERNAL
CANOPY HANDGRIPS
CANOPY SWITCHES
CANOPY LOCK
EMERGENCY
RELEASE HANDLE
EXTERNAL
CANOPY CONTROLS
PILOT’S CANOPY FAST-FIRE
JETTISON "T‘ HANDLE
# SOME AIRPLANES
I. PUSH BUTTON TO RELEASE HANDLE
^PULL'T'HAHOLfi CHIT & INCHES
TO JETTISON CANOPY
emergency entrance
CONTROL ON OTHER SIDE
CANOPY CONTROLS
O PILOT’S CANOPY
SWITCH AND
LOCK LEVER
*_ PILOT’S
CANOPY SLOW FIRE
JETTISON "T" HANDLE
Figure 1-40.
1-53
Section I
T.O. TF-89H-1
CANOPY.
The transparent canopy is operated by an electric mo¬
tor geared to a chain, and can be controlled normally
by any one of three switches: the pilot's, the radar
observer's, or the external switch. The canopy motor is
powered directly from the battery bus. In an emergency,
the canopy can be fast-jettisoned in flight by either
crewmember, slow-jettisoned on the ground by an ex¬
ternal emergency release, or slow-jettisoned on the
ground by either crewmember by a slow fire control
handle in each cockpit. The canopy travels fore and
aft on roller trucks and is sealed for pressurization by a
pneumatic seal that is automatically deflated and in¬
flated by movement of the canopy locks. The seal can
also be deflated by depressing the spring button on the
seal valve at the left of the pilot's left vertical console.
A brake on the actuating motor stops the canopy in any
position other than within the forward 10 inches of
travel, when the switch is released. When the canopy
closes to within approximately 10 inches of the closed
position, it trips a microswitch that deenergizes the
motor and allows the canopy to coast forward toward
the windshield. Just before the canopy strikes the
windshield (approximately 1 inch) another microswitch
energizes the actuating motor brake momentarily to
prevent the canopy from slamming into the windshield.
The canopy lock lever is then used to bring the canopy
to the locked position. A limit switch also brakes the
canopy motor to prevent the canopy from slamming
into the rear stops. Hydraulic dampers aid the actuating
motor brake in preventing the canopy from slamming
against the windshield or rear stops. This also provides
the needed braking action when the canopy is operated
manually and the actuating motor brake is inoperative.
On airplanes modified in accordance with T.O. 1F-S9-
600, the canopy push-pull circuit breaker has been
replaced with a toggle-type circuit breaker to facilitate
deactivation of the canopy system for ground operation.
WARNING
® When leaving the airplane, make certain that
no personal equipment, which could become
entangled with the seat armrests when the
canopy is closed or opened, is left in the cock¬
pit. Otherwise, the canopy may be accidently
jettisoned with attendant personal injury.
® When taxiing with canopy open, keep hands
clear of canopy track when applying brakes
as sudden brake application may cause the
canopy to slam forward.
Canopy Jettison System-
In an emergency, the canopy can be fast-jettisoned by
either crewmember by raising the ejection seat right
armrest, or by the pilot pulling out the canopy jettison
tf T" handle approximately 2 inches. The canopy can be
slow-jettisoned by the ground crew by pulling out the
external emergency release handle approximately 5
inches. The radar observer can slow-jettison the canopy
by using the emergency hydraulic pump handle to put
pressure against the cable attached to the external
canopy jettison lever and the canopy jettison initiator.
Either method releases compressed gas to the canopy
jettison cylinders. When the canopy is fast-jettisoned,
it is thrown clear of the airplane. When it is slow-
jettisoned, the canopy is slowly pushed above the cock¬
pit rails. From this position the canopy may be pushed
or lifted from the airplane. On airplanes modified in
accordance with T.O. 1F-89-586, both the pilot's and
radar observer's cockpits are equipped with an internal
canopy slow-fire jettison "T* handle. This enables
either the pilot or the radar observer to slow-jettison
the canopy by pulling the "T” handle. In the pilot's
cockpit the "T” handle is located on the left side below
the cockpit rail (figure 1-40). In the radar observer's
cockpit, the "T" handle is located below the main spar
on the left side (figure 1-40).
CANOPY EJECTOR PRESSURE GAGE.
The canopy ejector pressure gage (figure 4-6), located
on the radar observer's instrument panel, provides the
radar observer an accurate check of the canopy jettison
cylinder pressure.
PILOT'S CANOPY SWITCH,
A slide-type canopy switch (figure 1-40) on the handle
of the pilot's canopy lock lever is one of the three
spring-loaded switches that control canopy operation.
The switch positions are marked OPEN and CLOSE.
The switch is spring-loaded to an unmarked NEUTRAL
position. After the locks have been disengaged, the can¬
opy can be opened by holding the switch at OPEN until
the canopy has reached the desired position. When the
canopy is opened to its full limit of travel, a limit switch
operates a brake to keep the canopy from slamming
against the mechanical stops. To dose the canopy, the
switch is held at CLOSE until the canopy stops moving
and the lock lever is then pushed down to close and lock
1-54
T.O. 1F-89H-I
Section I
the canopy. The pilot's switch overrides the radar observ¬
er's switch, and the external switch overrides both cock¬
pit switches. Ail canopy switches operate on 28-volt dc
from the battery bus,
RADAR OBSERVER’S CANOPY SWITCH.
A spring-loaded canopy switch (figure 1-40) on the left
side of the radar observer’s cockpit is marked OPEN
and CLOSE and operates the canopy in the same man¬
ner as the pilot's canopy switch.
EXTERNAL CANOPY SWITCHES.
To permit electrical actuation of the canopy from out¬
side the cockpit, two battery-powered control switches
(figure 1-40) are located inside a key-locked access door
on the left side of the fuselage above the wing leading
edge. The two push-type switches are marked OPEN
and CLOSE. When either switch is held depressed, the
canopy moves in the desired direction until the switch
is released. The external canopy switches override the
pilot's and radar observer's canopy switches.
When opening the canopy with the external
canopy switch, use caution to prevent the for¬
ward corner of the canopy from striking the
operator's hand,
Nofre
If the canopy cannot be opened electrically,
open canopy manually.
CANOPY LOCK LEVERS AND INDICATOR LIGHT.
There are three canopy lock levers (figure 1-40): the
pilot's, near the floor at the left of the pilot's seat;
the radar observer's, on the left side of the cockpit; and
the external lever, just below the left cockpit rail in¬
side a key-locked external access door. Moving a lock
lever forward, when the canopy is within 1 inch of full
forward travel, fully closes and locks the canopy, and in¬
flates the canopy pressure seal. Pulling a lock lever
back releases the locks and a “canopy unlocked" 28-
volt d-c red indicator light next to the left windshield
defogging duct comes on.
Note
Prior to opening the canopy, place cabin air
switch to RAM & DUMP position to deflate
canopy seal.
The light goes out when the locks are engaged. The
external lever must be disengaged and pushed into its
dip for stowage.
PILOT’S CANOPY HANDGRIPS.
If 28-volt d-c electrical power is not available, the
canopy can be opened or closed manually. After release
of the canopy locks, the canopy is free to roll. Two
handgrips (figure 1-40) on the forward frame of the
canopy are for the pilot’s use in manual operation.
RADAR OBSERVER'S CANOPY HANDGRIPS.
The radar observer can move the canopy manually by
using U-shaped handgrips (figure 1-40) located on each
canopy rail.
EXTERNAL CANOPY HANDGRIPS.
Two external hinged handgrips (figure 1-40), one in
each side of the aft structure of the canopy, can be used
by personnel outside the cockpit to assist in manually
moving the canopy.
EXTERNAL EMERGENCY CANOPY
RELEASE HANDLE.
The canopy can be slow-jettisoned by an external
emergency release handle (figure 1-40) which is flush
with the fuselage skin just below the access door for
the external canopy switch. A button in the center of
the handle must be pressed in to release the handle.
Approximately 45 pounds of pull must be exerted to
break the safety wire on the jettison valve and a con¬
stant pull must be maintained until the canopy breaks
free and rises above the cockpit rails. When the handle
is pulled out approximately 5 inches and held, com¬
pressed gas flows through a restrictor to the actuating
cylinders and, in approximately 10 to 20 seconds, the
canopy will be pushed above the cockpit rails. From
this position it can be lifted or pushed from the air¬
plane.
The canopy should be jettisoned on the
ground only in an emergency. To prevent
accidental jettisoning of the canopy when the
1-55
Section I
T.O. TF-8SH-T
airplane is on the ground, safety pins must be
installed in the canopy jettison components
in both cockpits (as discussed in Ejection Seat
Ground Safety Pins, this section).
EJECTION SEAT RIGHT ARMREST.
The right armrest of either ejection seat (figure 1-41)
can be raised to fast-jettison the canopy. When either
crewmember raises his right armrest, compressed gas
under approximately 1800 psi flows to the actuating
cylinders, the canopy locks release, and the canopy is
thrown into the air.
% The canopy goes straight up when it is jet¬
tisoned. Lack of airstream may cause it to fall
back into the cockpit.
• If the canopy is to be jettisoned for reasons
other than ejection (such as a forced landing),
the pilot should not use the seat armrest, as
this will also cause his seat to bottom, thus
restricting vision. The canopy can be jetti¬
soned by the pilot without bottoming the seat
by pulling out the pilot's canopy jettison "T”
handle.
© Keep hands and arms clear of canopy lock
levers during canopy jettison. As the canopy is
jettisoned, the radar observer's lock lever will
rotate rapidly to the OPEN position and the
pilot's lock lever will snap to the up (OPEN)
position.
PILOT'S CANOPY FAST-JETTISON “T” HANDLE.
A “T” handle (figure 1-12), located on the pilot's right
vertical console, enables the pilot to fast-jettison the
canopy without using the ejection seat control. This
handle is linked by a cable to a gas initiator located on
the floor just forward of the right console. The cable
also is linked to a microswitch in the 2S-volt d-c circuit
to the canopy retention solenoids. Pulling the handle
out approximately 1 inch opens the microswitch and
interrupts the circuit to the canopy retention solenoids;
pulling the handle another inch (a total of approxi¬
mately 2 inches) draws the firing pin from the initiator
which in turn opens the shutoff valve to the canopy
jettison cylinders. The retraction mechanism for stowing
the radar observer's scope and console is then automati¬
cally actuated co the stowed position, and the jettison
cylinders release the canopy locks and throw the canopy
from the airplane. The pilot's canopy fast-jettison "T”
handle should be used for all emergencies, other than
ejection, requiring jettisoning of the canopy. To pre¬
vent inadvertent canopy jettisoning, a ground safety
pin is provided for the canopy jettison gas initiator.
This pin with its streamer is attached to the end of the
pilot's ejection seat ground pin streamer. On Group 5
airplanes, a canopy jettison "T” handle guard and a
large streamer are provided for additional safety,
EJECTION SEATS.
If the C-2A life raft is being carried, the A-5
seat cushion should not be left on the seat.
If both are used and it becomes necessary to
eject or crash land, severe spinal injury may
be caused by the excessive compressibility
of the combination of life raft and cushion.
If additional height in the seat is needed, a
solid filler block may be used in conjunction
with the life raft.
Note
When the seat cushion is not used, the Type
MD-1 contoured seat style survival kit con¬
tainer, stock number 2010-126602, with the
MA-1 contoured cushion, stock number 2010-
159100, should be used. The forward edge of
the packed kit should not be thicker than 7
inches (consult T.O. 14S1-3-51, “Base Assem¬
bly, Use and Maintenance of Sustenance Kits”
and T.O. 14S3-2-31, “One Man Life Raft,
Type PK-2, Used with Survival Kit Container,
Type MD-1"). The CA-2 one man life rafc kit
may be used if the MD-1 containers are not
available.
The pilot's and radar observer's stations are equipped
with catapult-type ejection seats (figure 1-41). A cata¬
pult aft of each seat contains an explosive charge that
supplies the propelling force for seat ejection. The
catapult is permanently safetied by two shear pins that
are sheared during firing by gas pressure from the
initiator. The headrest and footrests of each seat are
fixed. The pilot's seat is adjustable in combination
vertical-fore-and-aft directions. The radar observer's
seat is not adjustable. Controls for the ejection sequence
are the two armrests of each seat and the right hand¬
grip firing trigger. Movement of these controls actuates
a compressed air system that automatically lowers the
pilot's seat to the full down position, locks the shoulder
harness reel, fires the gas initiators which actuate the
components that jettison the canopy, stows the radar
1-56
T.O. 1 F-89H-I
Section I
H42(1)D
Figure f-41 (Sheet 1 of 7).
1-57
Section 1
T.O. 1F-89H-T
M
RIGHT ARMREST
LEFT ARMREST
RIGHT HANDGRIP AND FIRING TRIGGER
INERTIA REEL LOCK LEVER
SEAT ADJUSTMENT LEVER
H“42i2>D
1-58
figure J-4? fSheet 2 of 2 ).
T.O. 1F-89H-1
Section 1
PILOTS SEAT
EMERGENCY CANOPY
JETTISON "T" HANDLE
{Guarded, some airplanes)
SEAT SAFETY PINS
SAFETY BELT RELEASE
INITIATOR PIN
EMERGENCY CANOPY
JETTISON INITIATOR PIN
RADAR OBSERVER S SEAT
RIGHT armrest
GROUND SAFETY PIN
CANOPY JETTISON INITIATOR PIN
CATAPULT FIRING INITIATOR PIN
REMOVE BEFORE TAKEOFF
H-42t3)0
Figure 7-42.
1-59
Section ]
T.O. 1F-89H-1
observer's scope (stowed by raising right armrest of
either seat), and fires the catapult. As the seat is
ejected, anti "G" suit, oxygen hose, microphone, and
headset connections automatically disconnect at the
seat. For ejection, the canopy can be jettisoned by either
the pilot or radar observer, but seat ejection is con¬
trolled by the individual occupying the seat. An ejec¬
tion notification switch is installed on each crew¬
member's ejection seat. When either the pilot's or
radar observer's seat is ejected from the airplane, the
ejection notification switch automatically actuates the
emergency mode of the AN/APX-6 IFF system.
If time and conditions permit, the radar
observer rather than the pilot should jettison
the canopy. This will assure that the radar ob¬
server is in position for ejection and will have
no difficulty in reaching the ejection seat con¬
trols due to the wind blast or "G '* forces.
The safety belt releases automatically by means of gas
pressure from a delay initiator that is fired as the seat
is ejected, and allows approximately 2 seconds more for
the seat to clear the airplane before the safety belt is
released.
ARMRESTS.
The right and left armrests (figure 1-41) are not inter¬
connected and may be moved independently of each
other. Each armrest terminates in a loop-type hand¬
grip, the right handgrip containing the catapult firing
trigger. The pilot's and radar observer's armrests have
been painted gray and the handgrips orange-yellow
to focus attention on the actual ejection controls.
Each armrest is fitted with a jackknife-type brace that
is spring-loaded to assist the armrest into the full up
position, once the armrest is lifted free of its stowed
position. In normal flying position each armrest is
stowed in the full down position and held there by a
roller lock. Approximately 20 pounds upward pull is
required to pull the armrest through its first half inch
of travel. After that the assist braces snap the armrest
into the full up position where it is held in place by
spring tension and the overcenter action of the braces.
On either seat, raising the right armrest jettisons the
canopy, snaps the seat's catapult firing trigger up into
the ready position, and moves the radar scope into the
stowed position; in addition, on the pilot's seat, raising
the right armrest lowers the seat. Raising the left arm¬
rest locks the shoulder harness inertia reel.
# If canopy fails to jettison after raising the
right armrest, the pilot may pull the canopy
jettison "T" handle. If that system fails to
operate, raise the canopy locking lever and
move the canopy switch to OPEN. When the
canopy moves aft from the windshield frame,
the airstream will blow it from the fuselage.
If canopy fails to blow off when unlocked,
continue with normal ejection procedure and
eject through the canopy.
# Keep hands and arms dear of canopy levers
during canopy jettison. As the canopy is jet¬
tisoned, the radar observer's lock lever will
rotate rapidly to the OPEN position, and the
pilot's lock lever will snap to the up (OPEN)
position,
CATAPULT FIRING TRIGGER.
The catapult firing trigger (figure 1-41), located in
the loop-type handgrip of the right armrest, is locked
in the stowed position when the armrest is down in
normal flying position. When the right armrest is
raised, the trigger lock releases and the trigger is
snapped up into ready position. Squeezing the trigger
pulls the initiator firing pin, and gas pressure sufficient
to shear the permanent safety pins drives the catapult
firing pin into the detonator to fire the seat catapult.
SEAT ADJUSTMENT LEVER,
A lever at the forward right corner of the pilot's seat
bucket (figure 1-41) controls locking pins in the seat
adjustment mechanism. The lever rotates up and aft to
retract the locking pins in the seat positioning struts
aft of the seat bucket. When the lever is in the horizon¬
tal position the seat is locked in place. When the lever is
rotated up approximately 15 degrees, the locking pins
are withdrawn and the seat may be adjusted upward or
downward by relieving or applying weight to the seat
bucket. The spring-loaded "A” frame beneath the seat
exerts a constant upward lift on the seat bucket of
approximately half the weight of a pilot.
SAFETY BELT AUTOMATIC RELEASE,
The primary purpose of the safety belt automatic re*
lease (figure 1-44), particularly when used with an
automatic-opening aneroid-type parachute, is to extend
the maximum and minimum altitudes at which success¬
ful escape can be made using the ejection seat. In a
high altitude ejection (above 15,000 feet), the auto¬
matic system delays deployment of the parachute until
an altitude is reached where sufficient oxygen is
available to permit a safe parachute descent and air
1-60
TO* 1F-89H-1
Section I
RETAIN HOOK IN
THIS RING
AFTER TAKEOFF
ENGAGE HOOK
ON
AND LANDING
TO LAP
ARMING BALL
Figure F-43,
density is great enough to reduce parachute opening
shock* In a low altitude ejection, use of the automatic
system greatly reduces the overall time required for
separation from the seat and deployment of the para¬
chute, and consequently reduces the altitude required
for safe ejection* The various types of safety belt auto¬
matic releases have been thoroughly tested and are com¬
pletely reliable* Under no circumstances should the
automatic belt be manually opened before ejection,
regardless of altitude. Human reaction cannot possibly
beat the automatic operation of the release in opening
rhe safety belt and arming the parachute, particularly
under the stresses imposed by escape. The escape opera¬
tion using the automatic release is not only faster, since
it opens 2 seconds after ejection, but also protects the
crewmember from severe injury at high speeds* Because
the deceleration of a crewmember alone is considerably
greater than that of the crewmember and seat together,
immediate separation would result if the belt were man¬
ually opened just before ejection* This would not only
cause greater fI G*' forces during deceleration, but could
result in the parachute pack being blown open* The high
opening shock of the parachute under these eircum*
stances could cause fatal injuries. Currently, three types
of safety belt automatic releases are in general use, the
MA-1, the MA-2, -3, and -4, and the MA-5 and -6,
(See figure 1-44*) Any of these various types may be
found in the airplane. All three releases are designed
to be locked and opened manually under normal usage,
much the same as the standard manual safety belt, ex¬
cept that on the MA-1 through MA-4 models, a key
that is attached to the parachute lanyard must be in¬
serted into the release before it can be manually locked
to ensure that the crewmember does not overlook the
attachment of his parachute lanyard to the release*
(If an automatic parachute is not used, the key attached
to the release is used.) When the release is manually
opened, the key drops out of the release to prevent
inadvertently dumping the parachute. On the MA-5 and
-6 automatic releases, a ring on the end of the parachute
lanyard slips over the locking tongue of the release
mechanism; when the release is manually opened, the
ring slips free. However, on all three versions of the
automatic release, the key (or ring) remains attached
to the mechanism when the release is forced apart by
gas pressure following an ejection, thus actuating the
1-61
Section I
t-VW HO £'VK [MIX 9 VM HO S VW 3HAX
T-YW 3dAX
T.O. 1F-89H-1
Section I
parachute mechanism when the crewmember separates
from his seat. Manual operation of the system can over¬
ride the automatic features at any time. For example,
it Is possible to manually open the safety belt even
though initiator action has started. The parachute auto¬
matic features may also be overridden by manual opera¬
tion even though the automatic parachute ripcord
release has been actuated.
© If the safety belt is opened manually, the para¬
chute ripcord must be pulled manually.
© Improperly attaching the shoulder harness
and safety belt tiedown straps to the automatic
belt may prevent separation from the ejection
seat after ejection. To make the attachment
correctly, first place the right and left shoul¬
der harness loops over the manual release end
of the swivel link; second, place the auto¬
matic parachute lanyard anchor over the man¬
ual release end of the swivel link; then, fasten
the safety belt by locking the manual release
lever.
© The M-4 or M-12 safety belt initiator ground
safety pin with warning streamer must be
removed prior to flight. If the pin is not re¬
moved, automatic uncoupling of the safety
belt will not occur if ejection becomes neces¬
sary. if pin is installed, maintenance personnel
should be consulted on the status of the ejec¬
tion system before occupying the seat,
LOW ALTITUDE “ONE AND ZERO 5 ’ EJECTION
SYSTEM.
A system incorporating a one-second safety belt delay
and a zero-second parachute delay (“one and zero”) is
provided (some airplanes) for ejection seat escape sys¬
tems to improve low altitude escape capability. This
system utilizes a detachable lanyard (figure 1-43) that
connects the parachute timer knob to the parachute
"D” ring. At very low altitudes and and at low air¬
speeds, the detachable lanyard must be connected to
provide for parachute actuation immediately after sepa¬
ration of the aircrew member from the ejection seat.
At higher altitudes and airspeeds, the detachable lan¬
yard must be disconnected from the “D” ring, to allow
the parachute rimer to actuate the parachute below the
critical parachute opening speed and below the para¬
chute timer altitude setting, A ring attached to the
parachute harness is provided for the stowage of the
lanyard hook when it is not connected to the parachute
"D” ring. The connecting (hookup) and disconnecting
(unhooking) of the detachable lanyard and the para¬
chute “D” ring must be done manually by each crew¬
member. Prior to takeoff, the static cord lanyard should
be hooked up and the minimum safe ejection altitude
determined. After takeoff, the lanyard must be un¬
hooked and stowed by the crewmember after passing
through the minimum safe ejection altitude for his
particular system. Before landing, each crewmember
must hook up lanyard prior to reaching the minimum
safe ejection altitude for his system. After landing, the
parachute may be removed from the airplane with the
lanyard in the hooked-up condition. The following
table should be used to determine minimum safe ejec¬
tion altitudes for takeoff and landing. The figures
presented in this table are conservative for climbs, opti¬
mistic for descending conditions and applicable to level
flight attitudes. The “one and zero” and "two and zero”
are used during takeoff and landing emergencies only,
and the data for these systems are applicable to an
airspeed range of 140 to 300 knots IAS. The following
table should be used only as a guide because even
though a minimum safe altitude has been determined
prior to takeoff, the actual decision as to when to eject
in an emergency will be influenced by such circum¬
stances as airspeed, control, and attitude, as well as
altitude.
If the detachable lanyard has been installed
before the one-second safety belt initiater, a
“ two and zero” system is temporarily provided
wherein higher minimum safe ejection alti¬
tudes must be observed (see following table).
For nonautomatic parachutes used with automatic
safety belts, lanyard, part number 67C6200, w r ill be
used. The minimum safe escape altitudes specified for
one or two-second safety belt and zero second para¬
chute settings apply when the lanyard is attached to the
rip cord and safety belt.
1-Second
2-Second
Automatic
Automatic
Lap Belt
Lap Belt
(Ml2 Initiator)
(M4 Initiator)
2-Second Parachute
(F-1A Timer), B-4 or 5
Pack, C-9 Canopy
350 FT
550 FT
2-Second Parachute
(F-1A Timer), B-5 Pack
Oil Canopy
400 FT
600 FT
1-Second Parachute
(F-1B Timer), B4 or 5
Pack, C-9 Canopy
200 FT
350 FT
T-63
Section l
T.O* 1F-89H-1
PYLON FUEL TANK
FUEL FILTER DE-ICING
ALCOHOL TANK
CANOPY JETTISON AIR
BOTTLE FILLER VALVE
PILOTS SEAT
BOTTOMING
AIR BOTTLE
NOSE GEAR BUNGEE FILLER VALVE
FIRE EXTINGUISHER AGENT BOTTLE
HYDRAULIC RESERVOIR
RIGHT MAIN FUEL TANK
RADOME ANTI-ICING FLUID TANK
RADOME ANTI-ICING NOZZLE
BATTERY i
EMERGENCY AIRBRAKE BOTTLE FILLER *
ENGINE OIL
TANK
(Each side.)
CANOPY JETTISON ATR
BOTTLE
SINGLE-POINT FUEL FILLER
(UNDER WING)
HYDRAULIC ACCUMULATOR AIR FILLER
HYDRAULIC RESERVOIR
F L U 0 & SPECIFICATIONS
FUEL SPECIFICATION
RECOMMENDED
ALTERNATE
»
»
ENGINE OIL SPECIFICATION
HYDRAULIC FLUID SPECIFICATION
RADOME ANTI-ICING FLUID By Volume
»
»
ALCOHOL SPECIFICATION
FIRE EXTINGUISHING AGENT
SPECIFICATION
OXY {; KN SP KClUCATK >N
H-43IDD
Figure 1-45,
1-64
SERVICING DIAGRAM
Section 1
T.O, 1F-S9H-1
1-Second Parachute
(F-1B Timer), B5 Pack, 250 FT
C-ll Canopy
0-Second Parachute
(Lanyard to "D” Ring), 100 FT
B4 or B5 Pack, C-9
Canopy
0-Second Parachute
(Lanyard to "D” Ring), 150 FT
B4 or B5 Pack, C-ll
Canopy
400 FT
200 FT
250 FT
EJECTION 5EAT GROUND SAFETY PINS*
Ground safety for the ejection seats, when the airplane
is on flight status, is achieved by a canopy fast-jettison
"T” handle guard in the front cockpit and two safety
pins, one m the radar observer’s right armrest, and one
in the pilot’s armrest. The pin in the radar observer’s
cockpit is attached to a large red streamer. The safety
pin and "T” handle guard in the pilot’s cockpit are at¬
tached to opposite ends of a large red streamer. These
pins and guard are to be removed after the safety belts
are fastened and must be replaced before the belts are
opened. They should remain m the cockpit at all times*
Ground safety for ejection seats, during maintenance
operation, is achieved by additional safety pins which
are installed in each gas initiator, four in the front
cockpit and three in the rear cockpit. The points to be
safetied in each cockpit are the canopy fast-jettison
valve initiator, the catapult firing initiator under the
right armrest, and the safety belt release initiator on
the left seat frame, aft of the backrest. In the front
cockpit, a fourth point is the emergency canopy jet¬
tison initiator located on the floor forward of the right
console* The large red streamers attached to these safety
pins are fastened together with snaps. See figure 1-41.
SHOULDER HARNESS INERTIA REEL LOCK LEVER.
A two-position LOCKED—UNLOCKED shoulder har¬
ness inertia reel lock lever (figure 1-41) is used to man¬
ually lock the shoulder harness reel or leave it free, sub¬
ject to the inertia lock* The lever is located on the left
side of each ejection seat. The lever is held in position
by a friction disk and may be moved by a firm pressure
forward to lock, or aft to unlock, the reel. When the
lever is in the UNLOCKED position, the reel harness
cable will extend to allow leaning forward in the cock¬
pit; however, the inertia reel will automatically lock
the shoulder harness tension cable when an impact force
of 2 to 3 t4 GV’ is encountered. When the reel is locked
in this manner, it will remain locked until the lever is
moved to the LOCKED position and then returned to
the UNLOCKED position. When the lever is in the
LOCKED position, the reel harness cable is manually
locked to prevent bending forward. The LOCKED
position provides an added safety precaution over and
above that of the automatic inertia-operated safety lock.
The reel will also lock automatically when the left arm¬
rest is raised prior to seat ejection*
AUXILIARY EQUIPMENT.
Section IV of this manual describes the following
auxiliary equipment and its operation: cabin air-con¬
ditioning system, canopy defogging system, anti-icing
systems, communication and associated electronic
equipment, lighting equipment, oxygen system, auto¬
pilot, single-point fueling system, and miscellaneous
equipment. Armament is described in T.O. 1F-89H-1A,
a confidential supplement to this publication.
T-66
TO. 1 F-89H-1
Section I!
TABLE OF CONTENTS
Preparation for Flight . *.*. 2-T
Preflight Check ... t , , , 2-1
Before Starting Engines .. 2-7
Starting Engines. 2-7
Engine Ground Operation. 2-9
Before Taxiing. 2-9
Taxiing ......... . .2-10
Before Takeoff , r t . 2-10
Takeoff .... 2-12
After Takeoff—Climb .. .2-13
Climb ... .2-14
Cruise. 2-15
Flight Characteristics . .. 2-15
Descent .... . .2-15
Before Landing .. 2-15
landing .. ,2-18
Go-Around .,., ... .2-19
Touch-and-Go Landings..2-19
After Landing., ... 2-21
Stopping Engines. .2-22
Before Leaving Airplane.. 2-22
Abbrevioted Checklist ................... r 2-25
Procedure steps in this section are followed by the
symbols P, RO, or P—RO in parentheses to indicate
whether the particular step is applicable to the pilot,
radar observer, or both crewmembers.
PREPARATION ¥QR FLIGHT*
FLIGHT RESTRICTIONS,
Refer to Section V, Operating Limitations, for restric¬
tions and limitations.
FLIGHT PLANNING.
Prepare a complete flight plan to determine the re¬
quired fuel, oil, oxygen, airspeed, power settings, and
other items for the proposed mission. Use the operating
data in Appendix I to assist you in planning.
TAKEOFF AND LANDING DATA CARDS.
Fill out the takeoff and landing data cards using the
operating data in Appendix I to assist you.
WEIGHT AND BALANCE,
L Check takeoff and anticipated landing gross
weights and balance.
2. Make sure the airplane has been s erviced and
that the required armament and special equipment are
loaded.
3. Refer to Section V for weight limitations.
4. Refer to Handbooks of Weight and Balance Data,
T.O. 1-IB-40 and T.O. 1F-89H-5 for detailed loading
information,
5- Check that the weight and balance clearance,
DD Form 365 F, is satisfactory.
PREFLIGHT CHECK*
BEFORE EXTERIOR INSPECTION.
Check DD Form 781 for the status of the airplane;
make sure that the airplane has been properly serviced.
EXTERIOR INSPECTION,
Conduct the exterior inspection as shown in figure 2-1.
2-1
Section II
T.O. 1F-89H-T
mam msncrm
When approaching the airplane, note the
general Overall appearance and then
check the following items:
LEFT FORWARD SIDE
^ I* Pitot lube, static vents, ami probe dear,
2. Hydraulie fluid and radomc anti-icing fluid levels
cheeked; caps secured.
3. Nose wheel tires for condition, inflation,
and slippage.
4. Nose wheel door condition.
5. Nose wheel strut extension (approximately 3
inches) ; ground lock removed.
6. Static ground contact.
7. Fire extinguishing agent and bungee air pressures.
8. Landing-taxi light condition.
9. Battery access door—remove.
10. Radar pressure gages—check gages for pressure
and crystals for proper color,
11. Engine screen pressure gages—check for pressure.
12. Brake accumulator gage—608—2500 psi.
13. Emergency airbrake pressure gage—1500 ? 50 psi.
14. Battery connected and secured.
15. All access doors secured.
16. Angle-of-attack computer probe cover removed;
check freedom of movement.
17. Radar nose condition; anti-icing nozzle clear.
38. Sequence valve transfer piston for condition
position (out), so that lauding gear and doo:
sequence properly,
39. Gear tiplock unlocked, and roller free.
48* Wing leading edge condition.
41. Underside of wing for condition, fuel and
hydraulic leaks, tiedown ring flush, and fuel t
vent outlets free of obstructions.
42. Single-point refueling cap secured: refueling
door locked.
43. Pylon tank for security.
44. Tank pylon for condition; pylon vent port de
18.
19.
20,
21 .
22 .
23.
24.
25.
26.
27.
28,
29.
US %30.
31.
32.
33.
34.
35.
36.
37.
RIGHT FORWARD SIDE
Power panel and electrical accessories access door;
open and check all circuit breakers IN,
All access doors secured; right main tank filler
cap secured.
Hydraulic fluid level checked; cap secured.
Pitot tube and static vents clear.
Cabin pressure regulator outlet clear.
Engine intake duct clear; screens and compressor
blades aligned and undamaged; cheek screws inside
intake and accessory' section for security; check
ground for foreign objects.
Evidence of fuel, oil, and hydraulic leaks.
Engine doors secured. Door lock bolt position
indicators—locked position (some airplanes).
Engine air intake iloors free; external engine inlet
screens installed.
Check oil quantity; oil filler cap and dip stick
colter pin secured.
Eleventh-stage compressor bleed port clear.
Engine door No. 3 airscoop clear; inside door No,
3 air*coop—check for chafed fuel line. Engine
door No. 4 airscoop clear.
RIGHT WING
Wheel chocks in place; ground lock removed.
Tire condition, inflation, and slippage.
Brake disk for condition, pucks for proper
clearance, and brake shuttle valve checked.
Jack Jug pointing straight downward.
Landing gear outboard door condition; strut
extension (approximately 6 inches between torque
arm pivot points), Check outboard door locking
arm for tension.
Wheel well lines for condition and leaks.
Inboard main gear door closed and locked.
Bungee air pressure.
52. Frangible rocket tube covers installed.
53. Tip tank access doors secured.
54. Position light condition.
55- Tip pod fin for security of attachment.
56. Tip tank vent and fuel dump port clear.
57. Aileron and wing flap for condition and hydraulic
leaks; aileron neutral, wing flap up. Speed brake
external ground lock removed.
60.
61.
62.
63.
RIGHT AFT FUSELAGE
Tailpipe, fuel manifold, and flam eh older condition.
Eyelids condition and position: closed, J35-35A engines*
open, J35-35 engines.
Afterburner blast plate condition.
Refrigerator air intake and exhaust clear.
Aft fuselage access doors secured.
Fuselage position light condition.
%
14
14
EMPENNAGE
64, General condition,
65, Drain ports for hydraulic leaks.
66, Position lights condition.
67, Access doors secured.
68- Rudder for approximate
neutral position,
LEFT AFT FUSELAGE
69. Fuselage position light condition.
70. Afterburner blast plate condition.
71. Tailpipe, fuel manifold, and
flameh ol d er cond i lion,
72. Eyelids condition and position; closed,
J35-35A engines; open, J35-35 engines,
73. Oxygen filler door secured.
H-45C1>£
2-2
Figure 2-L
Changed 13 February 1959
T.O. 1F-89H-T
Section II
LEFT WING
74* Aileron and wing 1 flap for condition and hydraulir
leaks; aileron neutral, wing flap up. Speed brake
external ground look removed*
75- Tip tank fuel dump port and vent clear*
76* Tip pod fin for security of attachment.
77* Position light condition.
78. Millie Pod Launcher Accumulator pressure-
check ; access doors secured.
79* Missile doors flush.
80* Frangible rocket tube covers installed*
81. Anti-icing overboard duct clear.
82. Wing access doors secured.
83* Underside of wing for condition, fuel and
hydraulic leaks, tiedown ring flush, and fuel tank
vent outlets free of obstructions*
84* Pylon tank fuel level and amount checked;
filler cap secured*
85* Pylon lank pressure release valve closed
(some tanks).
86. Pylon and lank for condition and leaks*
87. Pylon tank sway braces secured*
88* Pylon tank for security of attachment.
89. Tank pylon for condition; pylon vent port clear,
90. Wing leading edge condition*
91* Wheel chocks in place; ground lock removed*
92* Tire condition, inflation, and slippage,
93* Brake disk for condition, pucks for proper
clearance, and brake shuttle valve checked*
94. Jack lug pointing straight downward.
95. Landing gear outboard door condition; strut
extension (approximately 6 inches between
torque arm pivot points). Check outboard door
locking arm for tension.
96* Wheel well zincs for condition and leaks*
97. Inboard main gear door closed ami locked.
98. Bungee air pressure*
m
99.
100 *
■r io2.
103*
104*
103 *
106,
107*
Sequence valve transfer piston for condition and
position (out), so that landing gear and door will
sequence properly*
(rear up lock unlocked, and roller free,
LEFT SIDE
Eleventh-stage compressor blew! port clear.
Engine floor No* 3 airscoop clear; inside door No.
3 airscoop—check for chafed fuel line* Engine
door No* 4 airscoop clear.
Check oil quantity; oil filler cap and dip stick
cotter pin secured.
Engine air intake doors free; external engine inlet
screens installed.
Engine doors secured* Door lock bolt position
indicators—locked position (sonic airplanes)*
Engine intake duet clear; screens and compressor
blades aligned and undamaged; check screws
inside intake and accessory section for security;
check ground for foreign objects*
Evidence of fuel, oil , and hydraulic leaks*
UPPER WING AND FUSELAGE
108* General condition of surface*
1(19* Tip tanks for equal amounts of fuel, pressure
release valves flush, and caps secured.
110* All fuel filler caps secured.
Ill, Static source outlets and cooling scoops on top
of fuselage clear.
112. Fuselage position light condition.
113- Emergency flap reservoir filler cap secured
(left.wing).
114* Alcohol tank; check quantity and cap secured
(right wing),
115. Canopy seal and windshield condition*
116* Windshield wiper condition*
117, Electrical access doors Ecrured,
118- Canopy control door secured ami emergency
release handle slowed.
H-45tt)E
Changed 13 February 1959
2-3
The cockpit is entered from
the left side of the airplane forward
of the wing . Kick-in steps and
handgrips are on the left side of
fuselage and the engine air intake
dttcL The canopy is unlocked manually
and opened by the external
canopy switch inside an access door
above the wing leading edge.
Section IK
CAUTION
Locate external power unit as far from the
airplane as the power cable will permit, to
reduce the hazard of fire from exhaust gas or
hot components of the power unit.
On some airplanes, two lockbolt position in¬
dicators on each engine nacelle door are pro¬
vided to permit visual reference of their
position when doors are being locked* When
the small inspection door cover plates are
removed, a movable lockbolt position indi¬
cator and a stationary reference indicator will
be visible* These indicators must be aligned
within 1/32 inch when the lockbolt is in
locked position*
HAND¬
GRIPS
KICK- IN
STEPS
ENTRANCE
For the proper method of entering the cockpit, refer
to figure 2-2*
BEFORE ENTERING COCKPIT
Check (1500—2000
L Canopy ejection pressure
psi)* (P—RO)
2* Ejection seats—Check* (P—RO)
Armrests and trigger stowed; safety pins in¬
stalled; safety belt initiator ground safety pin
removed; seat air bottle pressure 1600—1800
psi; catapult file mark aligned*
Note
If the safety belt initiator ground safety pin
is installed, consult maintenance personnel
regarding the status of the ejection system
before occupying the ejection seat*
Figure 2-2
Changed 13 February 1959
TO, 1F-89H-?
Section M
3. Circuit breakers—IN, (P—RO)
4. Armament switches—Check, (P)
Safety control switch—SAFE; mode switch—
SNAKE; salvo selector switch'—ZERO.
5. Flashlight—Check operation. (P—RO)
INTERIOR CHECK.
Front Cockpit.
Note
A pilot's checklist (figure 1-11) is located
above the center pedestal.
If the C-2A life raft is being carried, the A-5
seat cushion should not be left on the seat.
If both are used and it becomes necessary to
eject or crash land, severe spinal injury may
result because of the excessive compressibility
of the combination of life raft and cushion.
If additional height in the seat is needed, a
solid filler block may be used in conjunction
with the life raft,
L Armament switches—SAFE; armament safety
plug—Install. (P)
2. Safety belt and shoulder harness—Fasten; static
cord lanyard and automatic-opening parachute lanyard
—Connected; inertia reel operation—Check. (P—RO)
The M -4 or M-12 safety belt initiator ground
safety pin with warning streamer must be
removed prior to flight. If the pin is not re¬
moved, automatic uncoupling of the safety
belt will not occur if ejection becomes neces¬
sary. If pin is installed, maintenance personnel
should be consulted on the status of the
ejection system before occupying the seat.
3- Rudder pedals—Adjust, (P)
4. Battery switch—OFF. (P)
5. Throttles—Closed. (P)
6. 28-volt d-c external power—Connected (on right
intake duct). (P)
Note
® If more than 15 minutes are to elapse be¬
tween supplying power to the 28-volt d-c
bus and starting or operating engines above
idle rpm, place afterburner control switch
(circuit breaker) at OFF (unmarked) and
leave it OFF until just before starting engines.
This will deenergize the eyelid and altitude
idle bleed actuator solenoids, thus preventing
them from being damaged by overheating.
@ Check operation of all press-to-cest lights on
each control or indicator panel as the panel is
checked.
7. 115/200-voIt three-phase a-c external power—-
Connected. (P)
8. Exciter control switch—CLOSE momentarily. (P)
9. Alternator breaker control switch—TRIP mo¬
mentarily. (P)
10. Three-phase inverter switch—MAIN. (P)
11. Single-phase inverter switch—Check EMER¬
GENCY and NORMAL (leave on NORMAL). (P)
12. Alternator external power switch—CLOSE mo¬
mentarily. (RO)
13. Left console circuit breakers—IN. (P)
14. Emergency airbrake valve handle—OFF. (P)
15. Sideslip stability augmenter power switch—PWR
ON; rudder trim switch-—AUTO TRIM; electrical
rudder trim knob—Safety wired at center position, (P)
16. Single-point fueling light—OUT. (P)
17. Fuel control panel and fuel gages—Check. (P)
Crossfeed switch—CLOSED; fuel selector
switches—-ALL TANKS; system circuit break¬
ers—IN. Move fuel gage selector switch to each
position and note readings of right and left
indicators. Leave fuel gage selector at TIP so
that tip tank feeding can be checked while
taxiing out for takeoff.
After positioning the selector switch at any
position, allow at least 3 seconds to elapse
before selecting another position. This will
preclude any possibility of the affected fuel
system motor valves being reversed in mid-
cycle, which will cause shorter valve life.
18. Wing flap lever—TAKE-OFF. (P)
19- Left hydraulic system supplemental pump—
Check. (P)
Depress nose wheel steering button and watch
left hydraulic system pressure gage for pressure
buildup to 2500 psi.
% When a demand is made on the supplemental
pump by operation of any left hydraulic sys¬
tem control, the supplemental pump must not
be in operation for a period of time greater
than 6 minutes, followed by a rest period of
IS minutes,
® When no demand is made on it by operation
of any left hydraulic system control, the sup¬
plemental pump should not be in operation
for more than 30 minutes.
Changed 13 February 1959
2-5
Section II
TO. 1F-S9H-1
20. Speed brake—Check operation; leave closed. (P)
21. Operate all flight controls simultaneously. (P)
Visually check control surface operation.
22. Aileron and elevator trim switch—Check. (P)
Move the switch full travel to left, right, fore
and aft positions to make sure that the switch
automatically returns to NEUTRAL when re¬
leased. If the switch sticks in any of the posi¬
tions, enter this fact with a red cross on
DD Form 781 and do not fly the airplane.
During the check, stick force should be exerted
against the trim to assure that the trim can be
overpowered. Return the elevator trim to the
TAKE-OFF position when check is completed.
Check control grip for security.
:: CAUTION )i
In checking the control stick grip do not twist
the grip as such action may cause the grip to
become less secure.
23- Nose wheel steering button—Release. (P)
24. Left hydraulic system supplemental pump pres¬
sure switch—Check. (P)
Pump wheel brakes through several cycles to
drop brake accumulator pressure to between
1100—800 psi. Supplemental pump should come
on and brake accumulator pressure should start
to rise to approximately 2100 to 2350 psi.
L
CAUTION
**#**#**»#+#**##«*$
~f
• When a demand is made on supplemental
pump by operation of any left hydraulic sys¬
tem control, the supplemental pump must not
be in operation for a period of time greater
than 6 minutes, followed by a rest period of
15 minutes.
• When no demand is made on the supplemen¬
tal pump by operation of any left hydraulic
system control, the supplemental pump should
not be in operation for more than 30 minutes.
25. Position light switches—As required. (P)
26. Landing gear warning horn reset button—Press.
(P)
Landing gear lever light should come on.
27. Cabin temperature switch—AUTO. (P)
28. Cabin temperature rheostat—As required. (P)
29- Landing gear lever—Check DOWN. (P)
Check gear position indicators for safe gear in¬
dication. Emergency landing gear handle—
Check fN (stowed position).
30. Canopy seal button—Released, (P)
31- Landing-taxi light switches—As required. (P)
Check operation of both the landing and taxi
lamp beams after extending the light.
32, Windshield de-ice and defog knob—NORMAL.
<P>
33* Windshield wiper switch—OFF. (P)
34. Windshield wiper speed rheostat—INC, (P)
35. Anti-ice switches—OFF. (P)
36. Engine screen switch—NORMAL. (P)
37. Pitot heat switch—Check. (P)
Turn pitot heat switch ON and check opera¬
tion with crew chief. Leave on if necessary.
38. Canopy locking lever—UP, (P)
39- Cabin air switch—PRESSURE. (P)
40. Cabin differential pressure switch—5.00 PSL (P)
41. Attitude indicator warning flag—Retracted, (P)
42. Flight computer—Check. (P)
Flight computer selector switch—FLIGHT
INST; altitude switch—OFF; perform opera¬
tional check of flight computer (see Section
IV).
43- Directional indicator slaving cutout switch—IN
(P)
44. Altimeter and clock—Set. (P)
Cross-check with radar observer.
45. Parking brakes—Set. (P)
46. Canopy defog knob—IN. (P)
47. Fire and overheat warning circuits—Check opera¬
tion. (P)
Hold detector test switch to L & R FIRE CKT
1 and L OVERHEAT; left and right fire warn¬
ing 1 ights a nd left overhca t wa rn ing 1 ight
should come on within 2 to 10 seconds. Hold
to L & R FIRE CKT 2 and R OVERHEAT; left
and right fire warning lights and right over¬
heat warning light should come on.
48- Emergency signal button and light—Check.
(P—RO)
49- Starting power switch—NORMAL. For emer¬
gency start—EMER (see caution under step 57). (P)
50. Canopy jettison 'T Jf handle—IN (stowed posi¬
tion). (P)
51. Thunderstorm light rheostat knob—OFF. (P)
52. Interior and instrument lighting rheostats—As
required. (P—RO)
2-6
T.O. 1F-89H-1
Section II
53, Communication equipment—Check operation,
(P—RO)
Canopy must be closed to check the ARN-6 and
ARN-14, Radio compass—Check all positions
and set to desired frequency; UHF command
radio—Check all channels; YHF navigation set
—Check and set to desired frequency; inter¬
phone—Check operation,
54, Oxygen equipment—Check operation, (P—RO)
Pressure gage—400 to 450 psi maximum; warn¬
ing light switch—OFF; oxygen regulator diluter
lever—NORMAL OXYGEN; oxygen regulator
supply lever—ON; test system operation. (Re¬
fer to Oxygen System Preflight Check* Section
IV, for detailed information.)
55* Autopilot switches—OFF; turn knob—Centered.
(P)
56. IFF switch—OFF, (P)
operated under conditions of possible carbon monoxide
contamination, such as runup or taxiing directly be¬
hind another airplane, or during runup with the tail
into the wind, use the following procedure before
starting engines; Put on oxygen mask, connect tube to
oxygen regulator, and place diluter lever at 100%
oxygen. After contamination is no longer suspected,
place the diluter lever at NORMAL OXYGEN.
• The oxygen diluter lever must be returned to
NORMAL OXYGEN as soon as possible. Use
of 100 percent oxygen could deplete the sup¬
ply before the end of the mission.
57. Generator switches—ON, (P)
CAUTION !;
During emergency starts, one of the following
procedures must be used. If generator switches
are normally left in the OFF position they
must be turned ON (following engine start)
in the following order: left, right No. 2 and
right No. 1. If generator switches are nor¬
mally left in the ON position, the left gener¬
ator switch only must be turned OFF, then
turned ON after engines are started. Using
other than the above procedures may result in
the loss of the secondary bus and 2500 VA in¬
verter, or the tripping of the bus-tie relay cir¬
cuit breaker due to a current overload of the
left generator during right engine start (ex¬
ternal power connected). In either case the
right No. 2 generator should be turned ON
second, never first or third.
0 Before starting engines, make sure danger
areas (figure 2-3) fore and aft of the engines
are clear of personnel, airplanes, and vehicles.
Suction at the intake ducts is sufficient to kill
or seriously injure personnel if pulled against
or drawn into the ducts. Danger aft of the
engines is created by the high exhaust tem¬
perature and blast from the tailpipes.
0 To reduce foreign object damage to the en¬
gines, external engine and side door air inlet
screens will be installed for taxiing to or
from takeoff and landing areas and for
ground operations. The engines should be at
idle rpm or stopped during installation or
removal of screens as a safeguard to ground
crews. Personnel installing or removing the
screens shall approach from a 90-degree angle
and to the rear of the inlet duct opening. One
man shall stand at the wing tip of the airplane
to signal the pilot or operator in case of
accident.
58. Right console circuit breakers—IN. (P)
59- Right vertical panel circuit breakers—IN. (P)
60. Make sure all required navigational publications
are aboard. (P—RO)
BEFORE STARTING ENGINES.
Whenever possible, start and run up engines on a
concrete surface to prevent dirt and foreign objects
from entering the compressors and damaging the en¬
gines. Avoid runup on macadam pavement; high ex¬
haust temperatures may cause serious damage to the
pavement aft of the airplane. If the airplane is to be
CAUTION i;
ft#*################* t
Starting an engine by using the blast pro¬
duced by another aircraft or engine is pro¬
hibited. This method of starting engines
forces foreign objects into the intake of the
engine compressor section and results in en¬
gine failure.
STARTING ENGINES.
Start the left engine first, to supply hydraulic pressure
to the brake accumulator.
2-7
Section IE
T.O. 1F-89H-1
MMgjgJ
Am
°C)
MPH
34 MPH
I
90“F(32°C)
DISTANCE IN FEET
ENGINE AT MILfTARY
POWER 1100% RPM 1
(No Afterburner)
ENGINE AT TAXI
POWER <70^ RPM t
EXHAUST
temperature
EXHAUST
VELOCITY
EXHAUST
VELOCITY
EXHAUST
TEMPERATURE
500°F (260°C)
T 300° F fl49°<
I I
460 MPH 255 MPH
225 MPH 109 MPH
T
280°F(138°C)
165°F(74"C;
150
51 MPH
I
110°F{43°C)
22 MPH
T
82°F(28“C)
engine at maximum
POWER I 100“o RPM I
I EXHAUST
t VELOCITY
(If ith Afterburner)j exhaust
TEMPERATURE
.
685 MPH 340 MPH
T T
(316°C)
180 MPH
340°F ^(171 °C)
MPH
75 MPH
T
140°F (60°C)
' ■ NOTE
STANDARD DAY TEMPERATURE OF 60“ IS INCLUDED .]
IN THE ABOVE EXHAUST TEMPERATURES.
Figure 2-3,
lift moms.
L Fire guard posted, (P)
2. Throttles—CLOSED, (P)
3- Fuel selector switches—ALL TANKS; wing tank
switches—ON; tip tank switches—ON; pylon rank
switches—ON (if pylon fuel is carried), (P)
4. Crossfeed switch—CLOSED, (P)
5* Starter switch—START momentarily, (P)
Check for rise in oil pressure. If there is no in¬
dication of oil pressure immediately after start¬
ing* shut down engine and investigate.
6 . Throttle—IDLE when engine reaches 8 to 10%
rpm. (P)
The starter circuit should automatically discon¬
nect when load drawn by starter drops to 200
amperes (approximately 26% rpm). If ignition
does not occur within 5 seconds after moving
throttle to IDLE, close throttle and place starter
switch momentarily at STOP. Do not operate
the starter continuously for more than 1 min¬
ute. A second start may be attempted as soon as
the engine stops rotating. A 3-minute interval
must elapse after the second starting attempt
and a 30-minute interval must elapse between
each series of three starting attempts.
2-3
T.O. 1F-89H*!
Section II
The starter is limited to three starts of 1-min¬
ute duration each; if more than three starts
are required, allow starter to cool fot 30 min¬
utes before using again.
7. Exhaust gas temperature and rpm—Stabilized at
idle (49 to 51% rpm) after ignition. (P)
A hot start is a start during which the exhaust
gas temperature exceeds 915°C on J35-35 en¬
gines and 900X on J35-35A engines. After
any hot start during which the temperature
reaches 1000°C or five hot starts during which
the temperature is less than 1000°C, a "hot
section" inspection of the engine must be
accomplished. Exhaust gas temperatures be¬
tween 750°C and 9I5°C inclusive on J35-35
engines and 735 a C to 900°C on J35-35A en¬
gines are permissible for no more than 20
seconds. All hot starts must be entered in
DD Form 781.
8. Hydraulic system pressure gage—Check while
starting engine. (P)
When engine rpm is below 19% the pressure
should not exceed 400 psi; between 19% and
38% rpm purge valve should open; when en¬
gine rpm is above 38% the pressure should be
between 2800 and 3050 psi.
RIGHT ENGINE.
9- Right engine—Start as for left engine. (P)
10. External power—Disconnected, (P)
11. Battery switch—ON, (P)
12. Fuel pump warning lights—Off. <P)
13. Engine instruments—Check. (P)
Check for desired reading at idle rpm.
Note
When external power is disconnected, change¬
over to the airplane's 28-volc d-c system and
all three a-c systems is automatic.
ENGINE GROUND OPERATION.
No warmup period is necessary.
• During starting and accelerations, the maxi¬
mum allowable exhaust gas temperature is
915*C on J35-35 engines and900°Con J35-35A
engines. Exhaust gas temperatures between
750°C and 915°C inclusive on J35-35 engines
and 735°C to 900°C on J35-35A engines are
permissible for no more than 20 seconds.
• Do not exceed maximum rpm. If engine rpm
exceeds 104% momentarily or 103% stabilized,
with or without overtemperature, the engine
must be removed for overhaul. All overspeed¬
ing must be recorded in DD Form 781.
Note
See Section V for complete discussion on en¬
gine limitations.
BEFORE TAXIING.
Hold control stick back during ground tests,
VOLTAGE CHECK.
1- 28-volt generators—Check. (P)
With engines above 50% rpm, output of each
28-volt generator should be 27.5 volts; load-
meters should show 0.2 maximum permissible
difference,
2. Alternator—Check. (RO)
With the left engine above 60%, rpm, check the
output of the 115/200-volt alternator.
3. Both single-phase inverter buses and three-phase
inverter bus—Check output. (RO)
All three buses should read 115 volts with volt¬
meter selector switch at PWR INV PRI, PWR
1NV SEC, INST INV,
4. Three-phase inverter switch—SPARE. (P)
5. Three-phase inverter bus—Check output. (RO)
With voltmeter selector switch at INST INV,
voltmeter should read 115 volts.
6. Three-phase inverter switch—MAIN. (P)
7. IFF switch—STDBY. (P)
HYDRAULIC SYSTEM CHECK.
To check the left and right hydraulic flight control
systems individually, the left system must be checked
before starting the right engine.
1. Speed brakes—Check operation. (P)
2. Flight control surfaces—Check operation. (P)
Operate all flight control surfaces simultaneous¬
ly with both engines at idle rpm. Right hy¬
draulic system pressure should not drop below
1500 psi.
2-9
Section II
T.O. 1F-S9H-I
AUTOPILOT CHECK.
Perform rhe following autopilot check while taxiing
to save time and fuel,
L Autopilot power and autotrim switches—ON. (P)
Leave these switches ON for rhe duration of the
flight.
2. Turn knob—Check knob in DETENT position.
(P)
5. Engaging switch—ENGAGE. (P)
Move switch to ENGAGE after 1 1/2 to 2-
minute warmup. The switch should remain at
ENGAGE and the manual controls should re¬
sist movement.
4. Turn knob—Rotate clockwise and counterclock¬
wise; pitch knob—Rotate fore and aft, (P)
Stick should follow to right and left as turn
knob is moved; stick should follow fore and aft
as pitch knob is moved. Return knobs to DE¬
TENT position.
5. With nose wheel steering disengaged, yaw the
airplane to the right, then to the left with brakes. (P)
Left rudder pedal should move forward slightly
when airplane is yawed to the right; right rud¬
der pedal should move forward slightly when
airplane is yawed to the left.
6. Check force required to override autopilot. (P)
Operate the stick and the rudder pedals man¬
ually. Forces required to overpower the auto¬
pilot should not be excessive.
7. Autopilot emergency disconnect switch on control
stick—Squeeze. (P)
The engaging switch should return to the
DISENGAGE position and the controls should
be free.
TAXIING,
Maintain directional control with the steerable nose
wheel.
CAUTION
To reduce foreign object damage to the en¬
gines, external engine and side door air inlet
screens will be installed for taxiing to or
from takeoff and landing areas and for
ground operations.
L Ejection seat and canopy ground safety pins—
Removed. (P—RO)
2, Brake accumulator pressure—Check, (P)
3, Wheel chocks—Signal ground crew to remove,
(P>
4. Parking brakes—Release. (P)
5. Flight indicators—Check during taxiing. (P)
6, Perform autopilot check, (P)
7. Fuel gages—Check during taxiing for tip tank
feeding. (P)
Full tip tank fuel level indicates that tip tanks
are not feeding.
CAUTION
1
• Use of wheel brakes in addition to nose wheel
steering in turns will result in excessive stress
on the nose gear and excessive nose wheel tire
wear.
9 Nose wheel tires will be severely damaged if
maximum deflection turns are attempted at
rolling speeds in excess of 10 knots.
Use 70% to 75% rpm to start the airplane rolling from
a standing position and 50% to 55% rpm to keep it
rolling. If taxiing with left engine only, a higher rpm
is necessary. Maintain 60% to 70% rpm through turns
at low speeds. This requires a large clear area aft of
the tailpipes, A minimum of 115 feet of clear space
ahead of the airplane is required to make a turn safely,
starting from standstill. Minimize taxi time; flight
range is considerably decreased by high fuel consump¬
tion during taxiing. In addition, aircraft tires are not
designed to withstand extended durations of ground
rolling operations. Long taxi periods will build up
excessive temperatures and pressures in the tires, result¬
ing In decreased margin of safety and service life of
tires. Estimated fuel consumption for taxiing with
two engines operating is 30 to 70 pounds per minute;
therefore, 1 minute of taxi time costs from 3 to 8
nautical miles at long range cruising speed.
The engines must be at idle rpm or stopped
during installation or removal of screens as a
safeguard to ground crews.
BEFORE TAKEOFF.
PR£FLIGHT AIRPLANE CHECK.
After taxiing to takeoff position, complete the follow¬
ing check:
1. External engine and side door air inlet screens—
Removed. (P)
2-10
1\0. 1 F-89H-1
Section 11
Obtain clearance from ground crew that
screens have been removed. The engines must
be at idle rpm as a safeguard for the ground
crew.
2. Canopy—Closed and locked; warning light out.
(P)
3. flight controls—Check for free and correct move¬
ment. (P)
4 . Elevator trim—Check for TAKE-OFF setting, (P)
Be certain that airplane is trimmed properly
for takeoff. Excessive trim will cause danger¬
ous porpoising and possible stall.
PREFUGHT ENGINE CHECK*
Roll into takeoff position, center nose wheel, hold
brakes, and perform the following checks:
L Throttles—Full OPEN. (P)
Allow’ engine rpm to stabilize at 93 to 100%
rpm; observe exhaust gas temperature and check
other instruments for desired ranges.
Note
Acceleration from idle to 100% rpm takes
about 12 seconds for J35-35 engines and about
25 seconds for J35-35 A engines.
if CAUTION
Stabilized engine speeds greater than 103%
rpm or a momentary rpm of 104% or more
are prohibited, and engine must be removed
for overhaul if this overspeeding occurs. The
throttle must be reset if stabilized engine
speed exceeds 102% rpm.
5. Fuel selector switches—Check ALL TANKS. (P)
Use of ALL TANKS fuel selector position for
afterburner takeoff affords a greater margin
of fuel pressure for maintaining afterburner
operation than WING TANKS selector posi¬
tion because there is less flow resistance exist¬
ing in the fuel distribution lines from the
main tanks.
2. Fuel transfer—Check, with engines at military
power. (P)
Check fuel transfer by turning fuel selector
switch to MAIN. Low main tank fuel level w ill
indicate wing tanks not feeding. The aft CG
and/or main low level warning lights coming
on is further evidence that the wing tanks are
not feeding. Leave selector on MAIN.
3. Left afterburner—ON. (P)
Ignition will be indicated by thrust surge. Check
exhaust gas temperature and rpm stabilized,
4 . Right afterburner—ON. (P)
Check exhaust gas temperature and rpm sta¬
bilized.
6. Safety belt—Tighten; shoulder harness—Adjust to
fit snugly; inertia reel—UNLOCK; M L” shaped seat
safety pin—Remove, (P—RO)
7. Anti !t G” suit valve button—Press to check opera¬
tion, (P^RG)
3. Wing flap lever—TAKE-OFF. (P)
9. Speed brake lever—CLOSED. (P)
10. Attitude indicator—Set, (P)
II- Hydraulic flight control, brake accumulator, and
hydraulic reservoir pressure gages—Check. (P)
12, Autopilot powder and auto trim switches—ON. (P)
13, Check radar observer prepared for takeoff. (P)
1 4 , Engine screens—Extended (if any foreign objects
are likely to enter engine intake ducts). (P)
Note
Stabilization of rpm and exhaust gas tempera¬
ture takes approximately 3 to 4 seconds after
initiation of afterburning. The rise in exhaust
gas temperature and drop in rpm indicate
proper afterburner ignition. The subsequent
rise of rpm to normal indicates the opening
of the eyelids. Stabilization of exhaust gas
temperature is the final indication of eyelid
opening, afterburning, and airworthiness of
the engine.
5. Engine exhaust gas temperature and rpm—Check,
2-11
Section ti
T.0. 1F-89H-T
TAKEOFF PPOtFOOPF >
RELEASE WHEEL BRAKES,
KEEP NOSE WHEEL ON GROUND
UNTIL THE ABOVE APPLICABLE
AIRSPEED IS ATTAINED.
MAINTAIN DIRECTIONAL CONTROL WITH NOSE
WHEEL STEERING UNTIL RUDDER BECOMES
EFFECTIVE AT ABOUT 70 KNOTS IAS.
Note
© Determine normal exhaust gas temperature
(figure 5 - 2 ) for the existing runway tempera¬
ture prior ro takeoff. When engines have
accelerated to 100% rpm and before beginning
takeoff ground roll, check to ensure that
exhaust gas temperature is within limits. Be
sure to execute this check with the engine
anti-icing system deactivated, as the engine
anti-icing system, when actuated, may increase
exhaust gas temperature by as much as 20°C
(68°F). If the exhaust gas temperature is
abnormally low, sufficient thrust may not be
available for takeoff. Return to the line and
enter this information in DD Form 781.
# Ambient air temperature does not affect peak
temperature limits.
O If eyelids do not open, as indicated by the
afterburner warning lights remaining on
(some airplanes) and by excessive exhaust gas
temperature and drop in rpm, shut down
afterburner, retard throttles, and taxi back to
line.
• Except in cases of emergency, the engines
should never be shut down immediately after
afterburner shutdown. This practice tends to
permit accumulation of raw fuel in the after¬
burner, which may re ignite upon contact with
Figure 5-4.
hot engine parts. For normal operation it is
recommended that the engines be operated at
from idle to 70%, whichever rpm gives lowest
exhaust gas temperature, for at least 3 to 5
minutes after shutting down the afterburners.
This procedure will eliminate shroud segment
warpage, overheated bearings, and the possi¬
bility of raw fuel accumulating in the after¬
burners and igniting from hot engines.
TAKEOFF,
NORMAL TAKEOFF,
Note
The following procedure will produce the
results stated in the Takeoff Distance Chart
(figure A-6) in Appendix I.
When engines and afterburners are stabilized at 100%
rpm, proceed with takeoff as shown in figure 2-4.
See figure A-8 for refusal speed, and at checkpoint,
check airspeed.
WARNIH© \
Adhere closely to the recommended nose
wheel liftoff and takeoff airspeeds to assure
adequate lateral control and acceleration for
takeoff.
2-12
TO. 1F-39H-T
Section II
gradually EASE STICK back to lift nose
WHEEL ALLOWING AIRPLANE TO FLY ITSELF
OFF AT THE ABOVE APPLICABLE AIRSPEEDS:
AFTER TAKEOFF MAINTAIN APPROXIMATE
TAKEOFF ATTITUDE TO CLEAR A 5CXFOGT
HEIGHT AT 129 TO 153 KNOTS IAS,
DEPENDING ON GROSS WEIGHT
H47(2)t
Note
0 Takeoff with military power is possible, but
more distance is required. (See Takeoff Dis¬
tance Chart, figure A-6, for military power
takeoff distance.)
^ Sustained low-altitude operation at maximum
power can cause the rate of fuel consumption
from the main tanks to exceed the rate of
replenishment from the wing tanks. If the
aft eg warning light comes on under these
conditions, reduce power on the right engine
or increase altitude.
MINIMUM RUN TAKEOFF.
Strict adherence to takeoff procedure will result in
minimum takeoff ground run. For length of ground
run for various gross weights, see applicable Takeoff
Distance Chart (figure A-6).
OBSTACLE CLEARANCE TAKEOFF.
Follow normal takeoff procedure, using maximum
power. After attaining the 50-foot height IAS (see
After Takeoff—Climb, this section), maintain this IAS
until obstacles are cleared, then continue with normal
climb procedure.
CROSSWIND TAKEOFF.
Follow normal takeoff procedure with the following
exceptions: Use ailerons cautiously to maintain a wings
level attitude. Lift off at higher speeds than normal,
depending on wind velocity. Hold nose wheel on run¬
way until reaching takeoff speed to get maximum
benefit from nose wheel steering. This will greatly
reduce wheel braking. To determine component head¬
wind down the takeoff runway, and whether takeoff
is recommended under cross wind conditions at the pre¬
dicted minimum nose wheel liftoff speed, see Takeoff
and Landing Cross wind Chart (figure A-5).
r
CAUTION
Crosswind takeoff ground run distance can
be much greater than distances shown in the
Takeoff Distance Charts, depending on wind
velocity.
Note
Use of nose wheel steering will greatly facili¬
tate directional control during crosswind take-
off and minimize use of brakes.
AFTER TAKEOFF—CLIMB.
To gain altitude efficiently, first accelerate to the best
climb speed at constant altitude, then climb, maintain¬
ing the best climb airspeed according to the type of
climb desired. If a climb is started before reaching the
best climb airspeed, total time and fuel consumption
will be increased. The best power for climb depends
on the performance required. Maximum thrust, mili¬
tary thrust, or normal thrust may be used. Optimum
power settings for various performance requirements
are described in the following paragraphs.
2-13
Section II
TO. IF-89H-1
i. After takeoff, maintain approximate takeoff at¬
titude to clear a 50-foot height at airspeeds given in
applicable Takeoff Distance to Clear 50-Foot Obstacle
chart in Appendix. (P)
At takeoff airspeeds, aileron response may be
somewhat less than at higher airspeeds. Take¬
off airspeeds less than those recommended
will aggravate this condition.
4. After reaching a safe altitude, increase airspeed to
desired climbing speed, (P)
5. Static cord lanyard above minimum safe altitude
—Disconnect. (P—RO)
6. Fuel gages—Check, (P)
7. Fuel gage selector switch—ALL, (P)
8. Oxygen dilutee lever—NORMAL OXYGEN.
<P—RO)
Return diluter lever to NORMAL OXYGEN
as soon after takeoff as possible if takeoff was
made using 100% OXYGEN because of sus¬
pected carbon monoxide contamination of cock*
pit.
2. Landing gear lever—UP, when definitely air-
horne. (P)
I CAUTION
Landing gear and landing gear doors should
be up and locked and the light in the control
handle out before exceeding the structural
limit airspeed. Landing gear retraction at
speeds in excess of structural limit airspeeds
may result in partial gear retraction and pos¬
sible loss of or damage to the main inboard
landing gear doors. If ff G” forces or sideslips
are attempted during gear retraction, the
maximum airspeed at which the landing gear
will completely retract will be reduced.
Note
A priority valve in the hydraulic system gives
priority to all flight controls over landing
gear. Therefore, if the wing flaps are re¬
tracted before getting a safe up lock landing
gear indication, the gear movement will be
delayed until the flaps are up.
The oxygen diluter lever must be returned to
NORMAL OXYGEN as soon as possible. Use
of 100 percent oxygen could deplete the sup¬
ply before the end of the mission,
9, IFF switch—As required. (P)
CLIMB.
To gain altitude efficiently, first accelerate to the best
climb speed at constant altitude, then climb, main¬
taining the best climb airspeed according to the type
of climb desired. If a climb is started before reaching
the best climb airspeed, total time and fuel consump¬
tion will be increased. The best power for climb de¬
pends on the performance required. Maximum thrust,
military thrust, or norma! thrust may be used. Opti¬
mum power settings for various performance require¬
ments are described in the following paragraphs and
will produce the results stated in the applicable Appen¬
dix Climh Charts. During climb the following should
be accomplished at 5000 feet, 10,000 feet, and at level-
off altitudes:
3- Wing flap lever—UP after attaining a safe gear
and door UP indication and l60 knots IAS minimum
(170 knots IAS if full pylon tanks arc carried). (P)
CAUTION
Wing flaps must be fully retracted before
reaching structural limit airspeed to avoid
possibility of structural damage.
1. Oxygen—Check. (R—RO)
2. Altimeter and cabin altitude—Check for proper
operation, (P—RO)
3* Engine instruments—Check operation. (P)
4 . Wings and fuselage—Check. (P—RQ)
MAXIMUM RATE OF CLIMB.
To climb at the maximum rate (minimum time
climb), use maximum power and maintain airspeed
schedule shown in applicable Appendix Climb Charts,
2-14
T.O. 1F-89H-1
Section IE
MINIMUM FUEL CLIMB.
To climb using minimum fuel without regard to dis¬
tance gained, use military power at low altitudes and
maximum power above 20,000-foot pressure altitude.
Airspeeds shown in the applicable Appendix Climb
Charts are suitable for this type of climb,
MAXIMUM DISTANCE CLIMB.
To climb so that total distance covered, including
cruise distance, is greatest for fuel consumed, use mili¬
tary power and maintain the airspeed shown in the
applicable Appendix Climb Charts,
MINIMUM DISTANCE CLIMB.
Depending on gross weight and power, minimum dis¬
tance climb (maximum angle of climb) at low altitudes
may be obtained at the airspeeds shown in figure A-10.
Note
• During locked throttle climb, engine rpm
normally will not vary more than :±r2%.
• Minimum distance climb is not a maximum
rate of climb.
CRUI5E.
See Section VI and applicable Appendix charts for
cruise characteristics of the airplane.
FLIGHT CHARACTERISTICS.
See Section VI for flight characteristics of the air¬
plane.
DESCENT.
Any combination of power and speed brake position
may be used during descent if the airspeed limitations
in Section V are not exceeded. A normal descent pro¬
vides a compromise in fuel, time, and distance and is
ordinarily used during normal operation when loiter¬
ing or while awaiting landing clearance. The descent
is made at Mach 0.70 and idle power, maintaining the
airspeeds specified in the Descent charts (figure A-28),
With speed brakes fully open and engines at idle rpm,
descents up to 30,000 fpm can be made without exceed¬
ing 350 knots IAS. Use the following procedure in
making all descents:
L Throttles and speed brakes—As required. (P)
2. Windshield defrosting system—As required. (P)
3. Canopy defogging system—As required. (P—RO)
Operate windshield defrosting system as required. An¬
ticipate canopy fogging at low altitude and operate
defogging system accordingly. Speed brakes can be
used at any airspeed.
4. Altimeter—Set and cross-checked with radar ob¬
server prior to descent, (P)
BEFORE LANDING.
Before entering traffic pattern, airspeed may be varied
within wide limits with speed brakes. It is recom¬
mended that the pattern be entered at about 270 knots
IAS with speed brakes closed, using 85% rpm. If an
airspeed lower than 270 knots IAS is desired, open
speed brakes in preference to reducing power.
Note
• When power is stabilized at 85% rpm, ap¬
proximately 4 seconds are required to obtain
maximum power.
9 Because engine compressors are designed for
maximum efficiency at 100% rpm, compressor
efficiency will drop as rpm is decreased to
approximately 80% rpm. Therefore, if the en¬
gine is accelerated rapidly from 80% rpm to
maximum power, a compressor stall may re¬
sult. This is less likely to occur, however, at
85% or higher rpm since the compressor ef¬
ficiency increases quite rapidly with an in¬
crease in rpm.
1. Alert radar observer. (P)
2. Safety belt and shoulder harness—Tightened;
static cord lanyard—Connect prior to reaching mini¬
mum safe altitude; inertia reel lock lever—UNLOCK*
(P—RO)
3. Armament switches—OFF. (P)
Safety control switch—SAFE; mode switch—
SNAKE; salvo selector switch—ZERO,
4. Wing anti-icing system—OFF; engine anti-icing
system—As required. (P)
IjJARMIMy
Use extreme caution when using wing anti¬
icing during landing. Operation of the sys¬
tem causes a reduction in available thrust
which must be considered if a go-around is
necessary.
5. Windshield de-ice and defog knob—As required.
(P)
6. Landing light—As required. (P)
7. Brake accumulator and hydraulic pressure gages
—Check. (P)
8. Engine screens—Extended. (P)
Extend screens if any foreign objects are likely
to enter engine intake ducts.
9. Enter traffic pattern at 270 knots IAS, using 85%
rpm, (P)
10. Speed brake lever—OPEN. (P)
11. Airspeed 195 knots; speed brake lever—CLOSED.
(P)
12. Landing gear lever—DOWN; check gear down. (P)
2*15
Section II
TO. 1F-89H-!
NOTE
• Typical landing weight is based on a typical area
intercept mission , Weight includes fuel for
20-minute loiter at sea level plus 5 percent total
fuel and full armament.
© fncreoxe landing speed 2 knots above speed cited
on thh landing chart for each additional
1000 pounds increase in gross weight.
MAINTAIN A
MINIMUM OF
8S?o RFM UNTIL
LANDING IS
ASSURED.
H-4flfT)C
Figure 2-5.
TO. 1F-89H-1
Section II
AIRSPEED 195 knots; speed brake
LEVER—CLOSED: LANDING GEAR
LEVER-DOWN. CHECK GEAR POSITION
INDICATORS. VISUALLY CHECK MAIN
GEAR DOWN,
WHEN LANDING IS ASSURED.
RETARD THROTTLES TO IDLE.
WAj^Nim
• At higher gross weights, approach and
touchdown speeds must be increasetL
See landing speeds chart in appendix
for other weights and speeds.
H-46f2j'C
2-17
Section II
TO. 1F-89H-1
j| CAUTION ;|
Do not extend landing gear at airspeeds in
excess of the structural limit airspeed. After a
normal landing or during a two-engine go-
around the gear retraction cycle must be com¬
plete (gear door up and locked) before the
airplane exceeds the structural limit airspeed.
If practical, the structural limit airspeed
restriction should also be observed during
single-engine go-a round. In the event of simul¬
taneous actuation of landing gear, flaps, and
speed brakes, landing gear retraction time
will be lengthened. After rapid descent from
high altitude, allow for appreciably slower
landing gear and wing flap extension rates
caused by the low temperature of the hydrau¬
lic fluid.
13. Wing flap lever—-TAKE-OFF. (P)
14. Trim—Adjust as speed is reduced. (P)
15. Instruments—Check for desired ranges. (P)
16. Turn onto final at 170 knots IAS. (P)
17. Final approach: wing flap lever—DOWN; air¬
speed—Check. (P)
Speed brakes must be used with extreme cau¬
tion while on final approach. If speed brake
opening is increased rapidly, rapid decelera¬
tion may result in an excessive rate of descent
or stalling while still airborne.
18. Maintain 85% rpm until landing is assured. (P)
19. Maintain desired approach at 131 to 156 knots
IAS, depending on gross weight. (P)
20. When landing is assured, retard throttle to IDLE,
(P)
LANDING.
NORMAL t &NDING,
Mote
The following procedure will produce the
results stated in the applicable Landing Dis¬
tance Chart (figure A-29) in the Appendix.
For the landing procedure refer to figure 2-5. Aside
from the somewhat high stick force encountered dur¬
ing flareout, the airplane is easy to land. Tip and pylon
tanks must be emptied before landing to prevent ex¬
cessive loads in the tank attachment fittings. To avoid
hard landings (touchdown at coo high a rate of de¬
scent), do not open speed brakes fully until the air¬
plane touches down. With tail slightly down, touch
main gear down at applicable IAS given in Appendix
Landing Speed Chart. Rapid deceleration of the air¬
plane may result in stall while still airborne. If speed
brakes are closed just before touchdown, decreased
deceleration will result in a longer landing distance.
Open speed brake lever and set nose wheel down at
applicable airspeed.
Do not use nose wheel steering during nor¬
mal landing roll. Engaging the steering sys¬
tem while the rudder pedals are deflected
could result in an accident or structural
damage by causing the airplane to swerve
suddenly.
After nose wheel is down, apply wheel brakes inter¬
mittently ro avoid locking wheels. Only light brake
pedal pressures are required because braking action
is strong in comparison to the feel of the pedals. As
weight on the wheels increases with reduction in
speed, braking forces can be increased. Maximum brak¬
ing occurs just before tires begin to skid. Because of
the small tire tread and heavy weight of the airplane
the tires are easily skidded. Use nose wheel steering, as
required, for taxiing. See Section VII for added land¬
ing wheel brake information. Use nose wheel steering
as required for taxiing. Alternate use of the left and
right engines for single-engine taxiing will tend to
equalize taxi time on the two engines.
No fe
© See Landing Speeds Chart (figure A-30) for
landings at gross weights other than those
given in this section.
@ See Landing Distance Chan (figure A-29) for
total landing distance from a 50-foot height
using the normal landing procedure.
CROSSWIND LANDING.
Use normal landing procedure and correct for drift
as necessary on approach and landing. To determine
component headwind down the landing runway, and
whether landing is recommended under crosswind con¬
ditions at the predicted minimum nose wheel touch¬
down speed, see Takeoff and Landing Cross wind Chart
(figure A-5).
Note
Low aileron response will be experienced
below- 150 knots IAS.
2-18
T,G. 1F-89H-T
Section II
Do not select more than 1/3 full speed brake
opening prior to touchdown under crosswind
land ing trond it ions* Speed brake angles greater
than 1/3 full open will impair lateral control
as stall speed is approached*
HEAVY WEIGHT LANDING.
Anticipate a higher airspeed on the final approach and
also a greater ground speed and rolling distance with
increased gross weight. Begin braking at the applicable
speeds listed in the applicable Landing Distance Chart
{figure A-29),
MINIMUM RUN LANDING.
For a minimum ground run, normal landing pro¬
cedure is followed with one exception; the right en¬
gine is shut down immediately after three-wheel con¬
tact. The thrust eliminated by shutting down the idling
right engine will aid in reducing the landing roil.
Leave the wing flaps extended to take advantage of
aerodynamic braking on the landing roll. Exercise care
in brake application before the full weight of the air¬
plane is on the wheels, to avoid skidding.
WET OR ICY RUNWAY LANDING.
Anticipate a 20 to 30 percent longer landing roll {con¬
siderably greater for an icy runway landing) than
normal because of decreased braking friction. Use the
normal landing technique of full flaps with full speed
brakes immediately following touchdown and shut
down the right engine immediately after three-wheel
contact. Depend upon flap and speed brake drag for
initial deceleration, and apply wheel brakes cautiously
throughout the remainder of the landing roll to avoid
skidding. Leave wing flaps fully extended until after
turning off the runway. Open speed brakes after main
gear touches down and leave extended until after turn¬
ing off the runway,
GO-AROUND,
Because of slow engine and airplane acceleration,
make decision to go around as soon as possible. If a
landing cannot be completed, use the go-around pro¬
cedure shown below and in figure 2-6 as quickly as
possible:
1. Throttles—OPEN (afterburners on if necessary),
fP)
2. Speed brakes—CLOSED. (P)
3- Landing gear lever—UP (when definitely air¬
borne). fP)
4. Wing flap lever—As required. (P)
Gradually raise wing Haps as airspeed increases.
See figure 6-2 for applicable stalling speeds.
5. Clear traffic as soon as adequate airspeed is at¬
tained. (P)
Landing gear and landing gear doors should
be up and locked and the light in the control
handle out before exceeding the structural
limit airspeed.
TOUCH-AND-GO LANDINGS.
Touch-and-go landings should be made only when au¬
thorized or directed by the major command concerned.
Touch-and-go landings introduce a significant element
of danger because of the many rapid actions w'hich
must be executed while rolling on rhe runway at high
speed, or while flying in close proximity to the
ground. This type landing can be safely accomplished
with empty tip and pylon fuel tanks. Use caution in
performing the cockpit procedures while maintaining
directional control of the airplane. Use the following
procedures in performing touch-and-go landings:
Nofle
# Prior to making touch-and-go landings, per¬
form the Before Landing check.
® Maximum power should be used for all take¬
offs.
ON THE RUNWAY,
1. Throttles—Maximum power. (P)
2. Speed brakes—Closed. (P)
3. Wing flaps—Takeoff. (P)
4. Keep nose wheel on runway until nose wheel
liftoff speed is attained. (P)
5. Gradually ease stick back to lift nose wheel,
allowing airplane to fly itself off the runway at ap¬
plicable airspeed. {P)
AFTER TAKEOFF,
1. After takeoff maintain approximate takeoff atti¬
tude to clear a 50-foot obstacle at 129 to 154 knots IAS
depending on gross weight. Trim airplane to eliminate
excessive stick pressures, (P)
© It is important to adhere to applicable air¬
speed since stalling will be approached at a
lower airspeed, and takeoff distance will be
increased appreciably at a higher airspeed.
© At takeoff airspeed, aileron response may be
somewhat less than at higher airspeeds. Take¬
off airspeeds less than those recommended
will aggravate this condition.
2-19
Section H
TO. 1F-89H-1
CAUTION
Fuel required fur go-aroutirl is approximately S50
pounds with afterburning * find approximately
625 pounds without afterburning.
Do not exreed W5 knots IAS until landing gear
is up and limiting gear doors are closed.
NOTE
Landing gear retraction during go-around is slotrer
than normal because of the increased demands on the
hydraulic system by speed brake arari flap operation.
Landing gear retraction will be further slowed if
engine rpm drops below
H^9C
Figure 2-6.
2-20
TO, 1F-89H-1
Sect ion M
2. Landing lever—UP, when definitely airborne. (P)
Landing gear should be up and locked and the
light in the control handle out before exceed¬
ing 195 knots IAS. Landing gear retraction at
speeds in excess of 195 knots IAS may result
in partial gear retraction and possible loss of
or damage to the main gear landing gear
doors. If forces or sideslips are experi¬
enced during retraction the maximum air¬
speed at which the landing gear will com¬
pletely retract will be reduced.
3. Wing flap lever—UP. (P)
r
1
CAUTION |
Wing flaps must be fully retracted before
reaching structural limit airspeed to avoid
possibility of structural damage.
4. Fuel gages—Check. (P)
5. Throttles—As required to maintain desired alti¬
tude and airspeed, (P)
BEFORE LANDING (AFTER TOUCH-AND-GO].
When a series of touch-and-go landings are to be made,
reenter traffic pattern as Locally required. Enter the
traffic pattern using 85% rpm and maintaining ap¬
proximately 270 knots IAS with speed brakes closed.
If an airspeed lower than 270 knots IAS is desired,
open speed brakes in preference to reducing power.
Do not extend landing gear at airspeeds in excess of
195 knots IAS, After a normal landing or during a
two-engine go-around the gear retraction cycle must
be complete (gear door up and locked) before the air¬
plane exceeds 195 knots IAS. Prior to touch-and-go
landing perform Before Landing check.
Not©
After completion of last touch-and-go land¬
ing, perform the After Takeoff Climb of
After Landing check as applicable.
AFTER LANDING,
After nose wheel is down, apply wheel brakes inter¬
mittently to avoid locking wheels. Only light brake
pedal pressures are required because braking action is
strong in comparison to the feel of the pedals. As
weight on the wheels increases with reduction in
speed, braking forces can be increased. Maximum brak¬
ing occurs just before tires begin to skid. Because of
the small tire tread and heavy weight of the airplane
the tires are easily skidded. Use nose wheel steering as
required for taxiing. Alternate use of the left and right
engines for single-engine taxiing will tend to equalize
taxi time on the two engines.
CAUTION
Q When a demand is made on the supplemental
pump by operation of any left hydraulic sys¬
tem control, the supplemental pump must not
be in operation for a period of time greater
than 6 minutes, followed by a rest period of
15 minutes,
Q When no demand is made on the supple¬
mental pump by operation of any left hy¬
draulic system control, the supplemental
pump should not be in operation for more
than 30 minutes.
9 If carbon monoxide contamination is antici¬
pated during ground operation, oxygen
should be used w ith the diluter lever at 100%
OXYGEN,
9 Do not use nose wheel steering during a nor¬
mal landing. Engaging the steering system
while the rudder pedals are deflected could
result in an accident or structural damage by
causing the airplane to swerve suddenly.
CAl
CAUTION
9 Nose wheel tires will he severely damaged if
maximum deflection turns are attempted at
rolling speeds in excess of 10 knots.
• If the normal hydraulic brake pressure is lost,
release brake pedals, turn the emergency air¬
brake handle to ON, and operate the brake
pedals with caution . The emergency airbrake
system will supply enough pressure for three
complete brake applications.
2-21
Section II
TO. 1F-89H-1
Note
Adequate hydraulic pressure in the left sys¬
tem will he maintained during final approach
through actuation of the supplemental pump
by the landing gear lever switch. After touch¬
down, the pump will stop but will start again
as hrake accumulator pressure drops to be¬
tween 1100 and 800 psi when the airplane is
decelerated*
1. Turn off runway and come to a complete stop.
(P)
2. Safety pins—Insert in ejection sear and canopy
jettison mechanism. (P—RO)
3- Cabin air switch—RAM and DUMP (before open¬
ing canopy). (P)
4. Armament safety plug—Remove* (P)
5. With engines at idle have external engine screens
installed. (P)
6. Wing flap lever—UP. (P)
7. Speed brake lever—CLOSED. (P)
8. Trim—Reset to TAKE-OFF. (P)
9- Anti-icing, windshield de-ice and defog and pitot
heat switches—OFF. (P)
10. IFF—OFF. (P)
IL Taxi light—As required* (P)
STOPPING ENGINES.
To minimize the danger of explosion or fire
due to fuel vapor, park the airplane into the
wind when possible. Wait at least 15 minutes
after engine operation (flight or ground)
before going near the jet exhaust.
Note
The preceding procedure will eliminate pos¬
sible shroud ring segment warpage, over¬
heated bearings, and the possibility of raw
fuel accumulating in the afterburners and
igniting from hot engines*
3. Throttles—CLOSED* (P)
Move past IDLE stop to CLOSED by raising
fingerlifts* Throttle friction lever-—INCREASE.
6. Fuel selector switches—PUMPS OFF. (P)
7. All other switches—OFF except gene rator switches.
(P—RO)
BEFORE LEAVING AIRPLANE.
Surface control locks {except for speed brake locks)
are not necessary because of the irreversible hydraulic
control system.
1. Wheels—Chocked, and brakes released. (P)
2. All ground safety pins—Check installed. (P—RO)
3. Check that oxygen tube, radio cord, and personal
equipment are properly stow r ed. {P—RO)
• To prevent parachute from being opened in¬
advertently when wearing an automatic open¬
ing aneroid-type parachute that has a key at¬
tached to the aneroid arming lanyard, make
sure the key does not foul when leaving cock¬
pit.
@ When leaving airplane, make certain that no
personal equipment which could become en¬
tangled with the sear armrests when the can¬
opy is closed or opened is left in the cockpit*
Otherwise, the canopy may be accidentally
jettisoned with attendant personnel injury.
4 . Complete DD Form 781* (P)
1. Parking brakes—Set. (P)
2. Canopy—Open. (P)
3* Flight controls—Neutral* (P)
4. Engines—Run up before shutdown. (P)
If engines have been operating at normal rated
thrust or above (with or without afterburning)
for 5 minutes or more, either in flight or on
the ground, operate the engines at idle to 70%
rpm whichever rpm gives the lowest exhaust
gas temperature for at least 3 to 5 minutes
before shutting down, except in an emergency.
During flight operation, approach and taxi time
may be considered as part of this period.
F
CAUTION
To ensure inspection and maintenance of the
airplane, make appropriate entries in the
Form 781 covering any airplane limitations
that have been exceeded during the flight.
Entries must also be made when the airplane
has been exposed to unusual or excessive op¬
erations such as hard landings, excessive brak¬
ing action during aborted takeoffs, long and
fast landings, and long taxi runs at high
speeds.
2-22
T.O. IF-89H-1
Section II
5. Check pitot covers on; landing gear ground locks
installed, (P)
3N ]!
*#*W**tfs \
CAUTION
When leaving the airplane unattended, close
and lock the canopy. This inflates the canopy
seal, preventing moisture and dust from en¬
tering the cockpit.
Note
The following checklist is an abbreviated ver-
sion of the procedures presented in the ampli¬
fied checklists of Section 11. This abbreviated
checklist is arranged so you may remove it
from your flight manual and insert it into a
flip pad for convenient use. It is arranged so
that each action is in sequence with the am¬
plified procedure given in Section IL Presen¬
tation of the abbreviated checklist does not
imply that you need not read and thoroughly
understand the amplified version. To fly the
airplane safely and efficiently you must know
the reason why each step is performed and
why the steps occur in certain sequence.
2-23
1\0. 1F-89H-T
Section HI
Hf-3B
TABLE OF CONTENTS
Page
Engine Failure .. 3-1
Fire . , ... . . . ..3-13
Smoke and Fumes Elimination. 3-13
Ejection ..... . 3-13
Landing Emergencies. 3-17
Emergency Entrance. 3-20
Eimergency Exit on Ground. 3-20
Ditching .. 3-20
Oil System Failure...*. 3-21
Fuel Vent System Malfunction ............. 3-21
Fuel System Emergency Operation .. 3-22
Electrical System Emergency Operation. 3-24
Hydraulic System Emergency Operation ...... 3-27
Flight Control System Emergency Operation . . . 3-27
Sideslip Stability Augmenter Emergency
Operation .. 3-28
Wing Flap System Emergency Operation.3-28
Speed Brake System Emergency Operation .... 3-28
Landing Gear System Emergency Operation . . . 3-28
Brake System Emergency Operation. 3-30
Loss of Canopy. 3-30
Abbreviated Checklist. 3-31
Procedure steps in this section are followed
by the symbols P, RO, or P—RO in paren¬
theses to indicate whether the particular step
is applicable to the pilot, radar observer, or
both crewmembers.
ENGINE FAILURE.
SINGLE-ENGINE FLIGHT CHARACTERISTICS.
Single-engine directional flight control characteristics
are essentially the same as normal flight characteristics
because of the proximity of the thrust lines to the
centerline of the airplane. With one engine inoperative,
very little rudder trim is required. Thus, good control
is assured in the single-engine range. Minimum single¬
engine control speed is airspeed at stall. This airspeed
varies with gross weight, wing flap setting, speed brake
setting, and acceleration (such as that encountered in
banks and pull-ups). An airspeed of 160 knots IAS (170
knots IAS if pylon tanks are full) is a safe minimum for
all weights, all flap settings, all speed brake settings,
and moderate accelerations. See figure 3-1 for single¬
engine service ceilings. In single-engine flight where
only military power (100% rpm without afterburning)
is available on the operating engine, there are certain
airplane configurations in which level flight cannot
be maintained. At a typical takeoff gross weight of
42,000 pounds (pylon tanks empty or dropped), one
engine windmilling, flaps down 30 degrees or more,
and with the landing gear up or down , it is impossible
to maintain level flight. With the flaps up and the
landing gear up or down , level flight is possible; how¬
ever, until the landing gear is retracted or afterburning
initiated, performance will be marginal and any turns
or maneuvers may be accompanied by a loss of altitude.
SINGLE-ENGINE PROCEDURE*
Immediately after experiencing engine failure in flight
it is important to reduce drag to a minimum while
maintaining IAS and directional control while investi¬
gating for the cause of the engine failure. If the cause
of the malfunction cannot be determined, or if it is not
safe to continue operation, the procedure given below
should be followed for shutting down an engine in
flight.
1. Throttle (inoperative engine)-—CLOSED. (P)
2. Gear and flaps—Retract, if extended. (P)
3-1
Section 111
T.O. 1F-89H-1
The throttle for the inoperative engine should
he closed* If the throttle is left openj the
throttle controlled fuel shutoff valve will be
open allowing fuel to be metered through the
engine.
3. Engine fire selector switch for inoperative engine
—Raise guard and actuate switch. (P)
Mote
If the right engine fire selector switch is actu¬
ated and the right engine fuel selector switch
is at ALL TANKS, fuel from the right main
tank only will be available for crossfeed oper¬
ation. Wing tank fuel in the right system will
not be available until the fuel selector switch
is moved to WING TANKS.
4. Agent discharge switch—Actuate if necessary. (P)
I CAUTION jj
Do not actuate agent discharge switch unless
engine is on fire. This is a "one-shot” system,
and until the extinguisher bottle has been re¬
placed, there will be no further fire protec¬
tion available.
3. Generator switch(es) (inoperative engine)—OFF.
(P)
6. Inverter switches—As required, (P)
7. Unnecessary electrical equipment—Off, (P—RO)
8. Crossfeed switch—OPEN, (P)
9. Fuel selector switches—-ALL TANKS. (P)
If right engine is inoperative, do not operate
speed brakes unless left engine rpm is at least
85% or the supplemental pump is operating.
At lower rpm, the demand of speed brake
operation on the hydraulic system causes lim¬
ited aileron control unless supplemental pump
pressure is available.
SINGLE-ENGINE SERVICE CEILING
ALTITUDE HOT DAY
99 s F (37*C)AT SL
100% RPM
without AB
GROSS
WEIGHT
a
(“)
NOTE: All altitudes ore. pressure altitudes in feet
ALTITUDE STD DAY
59° F (15 °C) AT SL
96 % RPM
without AB
100 % RPM
with AB
100% RPM
without AB
96 a RPM
without AB.
100 ^ RPM
with AB
ff9|
mmm#*m
44,000
5900
9350
22,900
6400
19,800
40,000
9500
11,900
26,000
* 900
8900
23,000
36,000
13,000
14,800
29,700
* 7900
11,800
26,200
32,000
16,700
TB,300
32,900
13,900
14,850
29,600
DATA AS OF: 14 Aufimt [9^7 DATA BASIS: Hipht tvst
^ WITH POWER REDUCED TO PREVENT EXCEEDING EXHAUST GAS TEMPERATURE UMfT.
Figure 3-f.
3-2
TO* 1 F-89FM
Section 111
MINIMUM SAFI
SINGLE-ENGINE
SPIED
160
KNOTS
IAS
170
KNOTS
IAS
IF PYLON TANKS ARE FULL
ALL WEIGHTS FOR:
All flap settings All speed brake settings
HF131 i
10. Power on good engine—Readjust. (P)
11. Trim for straight and level flight. (P)
ENGINE FAILURE DURING TAKEOFF
(BEFORE LEAVING GROUND).
Takeoff Aborted.
If an engine fails before leaving the ground, continu¬
ing the takeoff depends upon length of runway, config¬
uration, gross weight, airspeed at time of failure, field
elevation, and ambient temperature. To help the pitot
make a decision, single-engine takeoff distances for
various gross weights, altitudes, and ambient tempera-
cures are shown in the Appendix, figure A-6. This
chart gives entire takeoff distance with one engine
operating at maximum power, and is to be used only
if an engine fails during the takeoff roll. If a decision
to stop is made, use the following procedure:
L Alert radar observer. (P)
2. Throttles (both engines)—CLOSED. (P)
3. Nose wheel steering button—As required. (P)
4 . Speed brake lever—OPEN. (P)
5. Wheel brakes—Apply (maximum braking occur:
at a point just before tires skid). (P)
If hydraulic pressure is insufficient for ade¬
quate braking, depress the nose wheel steer¬
ing button. This energizes the 2500-psi sup¬
plemental pump which will provide adequate
pressure for braking. If this should fail, use
the emergency airbrake system.
6. Emergency airbrake system—As required. (P) If
hydraulic pressure is insufficient for adequate braking,
use the emergency airhrake system.
7. Canopy—Jettison with canopy jettison “T" handle.
(P)
8. Inertia reel—Lock, (P —RO)
Note
All equipment should be set as required
before locking inertia reel, as some smaller
pilots may find it difficult to reach such
items as the canopy fast jettison "T” handle
after the inertia reel is locked.
9. Steer for smoothest terrain if remaining runway
is insufficient for stopping. (P)
10. If necessary, raise landing gear by depressing the
emergency override lever and simultaneously moving
the landing gear lever to UP. (P)
Note
@ Leave landing gear lever in DOWN position
if runway is equipped with Type MA-1A
runway overrun barrier and aircraft is modi¬
fied to contain the necessary arresting gear
equipment.
# If the left engine fails, depress the nose wheel
steering button. This will energize the left
hydraulic system supplemental pump which
in turn will supply adequate hydraulic pres¬
sure to all units normally supplied by the left
hydraulic system engine-driven pump.
Changed 13 February 1959
3-3
Section HI
T.O. 1F-89H-7
EMERGENCY OVERRIDE
OPERATION
Figure 3-2.
IL Engine fire selector switches—Raise guards and
actuate. (P)
12. Agent discharge switch—Actuate. (P)
13. Battery switch—OFF. (P)
1 4 . Generator switch es—OFF. {P)
15. When stopped—Abandon airplane, (P—RO)
ENGINE FAILURE DURING TAKEOFF
(AFTER LEAVING GROUND).
If an engine fails immediately after takeoff, lateral
and directional control of the airplane can be main¬
tained if airspeed remains above stalling speed, but
ability to maintain altitude or to climb depends upon
gross weight, airplane coo figuration, and air density.
Figure 3-3 shows the maximum gross weights at which
a 100 feet-per-minure rate of climb can be maintained
with landing gear down and flaps at takeoff for clean
configuration and pylon tank configuration. For en¬
gine failure immediately after takeoff, use the follow¬
ing applicable procedure:
Takeoff Continued.
L Warn radar observer of engine trouble. (P)
2. Landing gear and flaps—As required. (P)
Note
• If an immediate obstacle must be cleared, do
nor retract gear until obstacle is cleared. Re¬
traction of the gear creates considerable addi¬
tional drag. If the airplane is accelerating and
no immediate obstacle must be cleared, the
gear should be retracted.
• If the left engine fails, depress the nose wheel
steering button. This will energize the left
hydraulic system supplemental pump which
in turn will supply adequate hydraulic pres¬
sure to all units normally supplied by the left
hydraulic system engine-driven pump.
3. External stores—jettison. (P)
When the pylon tanks are jettisoned manually
(gravity drop), minor damage to the airplane
may result.
4. Tip tank dump button—Press. (P)
To completely dump full tip tanks will re¬
quire approximately 90 seconds; therefore,
the weight reduction will be gradual rather
than instantaneous.
5. Throttle (inoperative engine)—CLOSED. (P)
6. Engine fire selector switch for inoperative engine
—Raise guard and actuate. (P)
7. Agent discharge switch—As required. (P)
8. If obstacle must be cleared, hold airspeed at mini¬
mum safe value above stall to maintain best angle of
climb. (P)
9- After obstacle is cleared, allow airspeed to in¬
crease to 160 knots. (P)
10. Wing flap lever—UP, at 160 knots. (P)
Do not raise wing flaps below 160 knots IAS
because maximum lift will be reduced, pos¬
sibly below the minimum required to main¬
tain altitude.
11. Electrical equipment (nonessential)—Off. (P—
RO)
3-4
T.O. 1F-89H-1
Section Ml
12, Trim—As required* (P)
13* Cross feed switch—OPEN* (P)
14* Generator switch (es) for inoperative engine—
OFF. (P)
Continued Flight Impossible.
L Warn radar observer of impending forced land¬
ing. (P)
2* Lower nose to maintain flying speed. Prepare to
land straight ahead if possible. Alter course only to
miss obstacles* (P)
3. External stores—Jettison* (P)
© Do not jettison armament unless area is suit¬
able,
© Do not dump tip tank fuel as this will increase
fire hazard*
4. Landing gear lever—DOWN* (P)
GAR-1 missiles can be jettisoned only by fir¬
ing them in an unarmed condition* There¬
fore* make certain area ahead of airplane is
uninhabited*
5* Wing flaps—As required (if left engine or sup¬
plemental pump is operating). (P)
6* Speed brakes—As required* (P)
7. Throttles—CLOSED* (P)
Do not dump wing tip fuel as this will in¬
crease fire hazard*
Weight in pounds at which 100 feet per minute
Without pylon tanks rate of climb can be maintained with gear down ?
flaps in takeoff position, and maximum power .
FIELD ELEVATION AMBIENT TEMPERATURE
{Feet)
-10°C (+14 C F)
*10°C (*50°F)
*30°C (+86°F)
+50°C (+122°F)
5000
40,100
37,050
31,920
4000
41,620
38,200
33,140
3000
43,175
39,760
34,350
2000
43,175
41,200
35,640
28,800
1000
43,175
42,780
36,980
29,850
SEA LEVEL
43,175
43,175
38,300
30,900
DATA AS OF: 14 August 1957
DATA
BASIS: Flight Tent
H-528
Figure 3-3.
3-5
Section III
T.O. 1F-89H-1
8. Canopy—Jettison with canopy jettison "T” ban-
die. (P)
9* Inertia reel lock lever—LOCKED. (P—RO)
10. Engine fire selector switches—Praise guards and
actuate. (P)
11. Agent discharge switch—As required. (P)
12. Generator switches—OFF. (P)
13. Battery switch—OFF just before touchdown. (P)
Note
When the battery switch is placed at OFF,
the left hydraulic system supplemental pump
will be deenergized.
14. When stopped—Abandon airplane. (P—RO)
ENGINE FAILURE DURING FLIGHT (LEFT OR RIGHT
ENGINE).
If an engine fails during flight, investigate to deter¬
mine the cause before attempting an air restart. It is
recommended that the fuel system be checked first
for proper operation. If the failure is caused by im¬
proper fuel system operation and the condition is cor¬
rected, restart the engine. (See Restarting Engine in
Flight, this section.) If failure is caused by mechanical
breakdown, as may be indicated by engine instruments
or excessive vibration, the engine should be shut down.
See figure 3-1 for single-engine service ceilings and
applicable appendix charts for single-engine operating
data. For procedure on shutting down engine in flight,
see Single Engine Procedure, this section.
Note
If both engines fail and no restart is to be
attempted, turn battery switch to OFF to
conserve electrical power needed to operate
the left hydraulic system supplemental pump
for an emergency landing.
If both engines fail, battery switch must be
turned to ON again to operate supplemental
pump.
RESTARTING ENGINE IN FLIGHT,
For best starting conditions and wherever practical,
attempt air starts at 20,000 feet or below.
Hot®
Before a restart is attempted and the igniter
plugs are energized, fly the airplane in a nose
high attitude (5 to 10 degrees above the hori¬
zontal) to drain excess fuel from engine.
A normal air restart can be made if the engine rpm is
at least 12.5% and the airspeed is approximately 170 to
250 knots IAS, If both engines have failed, no attempt
should be made to restart both engines at the same
time. Battery power may be insufficient for simultane¬
ous ignition of both engines; therefore it is recom¬
mended that only one engine be started at a time.
Successful air starts after double flameout are depen¬
dent upon sufficient altitude and battery power. When
above 20,000 feet, conserve battery power while de¬
scending to 20,000 feet or below, by turning fuel selec¬
tor switches to OFF position, main power inverter to
OFF position, and all other unnecessary electrical
equipment off. Normally the left engine will be started
first, unless there are known reasons for a hazardous or
unlikely left engine restart. Place the fuel selector
switch, for the engine to be restarted, in a position
other than the position existing at the time of flame¬
out, provided there is sufficient fuel in the new selec¬
tion. This will cause relays and valves to recycle and
may clear up the difficulty. Place the crossfeed switch
at CLOSED and turn the power inverter to ON. Restart
the selected engine and when rpm and exhaust gas
temperature are stabilized, restart the other engine. If
the second engine fails to start, place the crossfeed
switch to OPEN and attempt another start. If a double
flameout is experienced at low altitude, the fuel selec¬
tion for the engine to be restarted first should be
changed, provided there is sufficient fuel remaining in
the new selection and time permits. The following pro¬
cedure should be used for all air starts:
s CAUTION j!
• Do not attempt to restart both engines at the
same time,
# Failure to windmill at least 12.5% rpm indi¬
cates damage to an engine. Under this condi¬
tion, do not attempt an air start.
L Throttle—CLOSED. (P)
2. Fire selector switch—Check OFF. (P)
3* Fuel selector switch—Change tank selection pro¬
vided there is fuel remaining in new selection. (P)
4. Cross feed switch—CLOSED. (P)
5* Power inverter—ON. (P)
6. Altitude start switch—ALTITUDE START mo¬
mentarily (for selected engine only). (P)
7. Throttle—IDLE (rpm and exhaust gas tempera¬
ture stabilized) then advance to desired rpm. (P)
3-6
T.O. 1F-89H-1
Section 111
ALTITUDE—FEET
HOLD THE FOLLOWING IAS
Figure 3-4.
8. If scarring is unsuccessful* attempt another start
at a lower altitude, In case of double fiameout, reduce
electrical load and attempt another start at lower
altitude. (P)
MAXIMUM GLIDE.
For the distance this airplane will glide, power off*
at various gross weights, refer to figure 3 - 4 . During
descent, the speed of the windmilling engines will
be high enough to supply power to the hydraulic
system for normal descent operation of the flight
controls, provided that engine rptn on either engine
does not drop below 10%. The supplemental pump
should be used to ensure adequate control for landing;
but to conserve battery power, the system should be
left off until final approach. This can he done by
turning off the battery switch before lowering the
landing gear, lowering the landing gear with the
emergency handle, and turning the battery switch on
again when turning onto the final approach.
CAUTION I
^ Supplemental pump must not be in opera¬
tion for more than 6 minutes* followed by a
rest period of 15 minutes, when a demand is
made on the pump by operation of any left
hydraulic system control.
# Supplemental pump should not be in opera¬
tion for more than 30 minutes when no de¬
mand is made on the pump by operation of
any left hydraulic system control.
3-7
Section Ilf
T.O. T F-89H-I
The downwind leg of the pattern should be extended
for a single-engine landing so that a lower than nor¬
mal approach angle will be flown, thus allowing the
use of higher engine rpm in case a go-around is neces¬
sary. Wing flaps are available with either or both en¬
gines inoperative* In the event of electrical failure, the
radar observer can normally maintain brake accumu¬
lator pressure by pumping the hydraulic handpump so
that the emergency airbrake system need not be used.
If it becomes necessary to use the airbrakes during the
landing roll, the pilot should apply the brakes care¬
fully since they are very sensitive and effective. Do
not pump the brakes because air is lost each time pedal
pressure is released*
The battery will supply power for the opera¬
tion of the supplemental hydraulic pump for
a very limited time only, even with the elec¬
trical load reduced to a minimum*
LANDING WITH ONE ENGINE INOPERATIVE.
If a landing with one engine is necessary, dump tip
tank fuel and drop pylon tanks. Approach the airport
at 250 knots IAS using no more than the following
engine rpm:
Ambient Temperature Engine RPM
50°C 93%
30°C 92%
I0°C 91%
If more than above power is required to
sustain level flight at 250 knots IAS, gross
weight must be reduced before landing; other¬
wise, reserve power may not be adequate to
main tain desired approach path after landing
gear and flaps are lowered.
Note
At airspeeds below 160 knots IAS, it may be
necessary to lose altitude in order to increase
airspeed* Bear this in mind if single-engine
landing becomes necessary and there is the
slightest chance that a go-around may be
necessary*
Do not extend flaps below the takeoff posi¬
tion. If flaps are extended below takeoff posi¬
tion, they must be raised to at least the takeoff
position in case of a go-around. Single engine
go-around with flaps in full down position is
impossible because level flight cannot be
maintained.
Right Engine Inoperative.
See Single-Engine Landing Pattern, figure 3-5.
1* Airspeed—Decelerate to 195 knots IAS on down¬
wind leg* (P)
2. Landing gear lever—DOWN. (P)
Note
Lowering the landing gear by the emergency
procedure will not affect subsequent normal
operation.
3. Wing flap lever—TAKE OFF. (P)
Do not extend flaps below the takeoff posi¬
tion, If flaps are extended below takeoff posi¬
tion, they must be raised to at least the takeoff
position in case of a go-around* Single engine
go-around with flaps in full down position is
impossible because level flight cannot be
maintained.
4. Airspeed—Stabilize at 180 knots IAS. (P)
5. Turn onto final and stabilize at 160 knots IAS.
Fly a lower than normal approach angle so that high
3-8
T,0, IF-89H-1
Section III
rpm can be used. Use of high rpm will reduce the
time needed to obtain maximum power should a go-
around be necessary. (P)
6. Do not reduce airspeed below 160 knots IAS until
landing is assured. (P)
7. Throttle—Retard to idle only when positive of
landing. (P)
8. Speed brakes—OPEN, after touchdown to reduce
ground roll. (P)
Left Engine Inoperative*
See Single-Engine Landing Pattern, figure 3-5,
L Airspeed—Reduce to 195 knots IAS on down-
wind leg, (P)
2. Landing gear lever—DOWN. (P)
3- Landing gear emergency release handle—PulL (P)
4. Landing gear position—Check, (P)
5. Emergency landing gear release handle—
STOWED. (P)
6. Wing flap lever—-TAKE OFF, (P)
7. Airspeed—Stabilize at 180 knots IAS, (P)
8. Turn onto final and maintain 160 knots IAS,
Fly a lower than normal approach angle so that high
rpm can be used. Use of high rpm will reduce the
time needed to obtain maximum power should a go-
around he necessary. (P)
9. Maintain 160 knots IAS until landing is assured.
(PI
10. I h rot tie—Retard to IDLE when positive of land¬
ing. Speed brakes—OPEN after touchdown. (P)
LANDING WITH BOTH ENGINES INOPERATIVE.
See Forced Landing, figure 3-6.
SINGLE-ENGINE GO-AROUND*
T he greater the speed when the decision is made to go
around, the shorter the go-around distance. If doubt
exists as to the landing, an immediate decision to go
around will save considerable distance. When the go-
around decision is made, complete the following steps
in the order given:
1. Throttle (on operating engine)—OPEN. (P)
2, Afterburner (on operating engine)—ON (above
90% rpm). (P)
3. Speed brake lever-—CLOSED. (P)
4, Wing flaps—20 degrees. (P)
• Single-engine go-around with flaps in full
down position must never be attempted be¬
cause level flight cannot be maintained.
@ During level flight accelerations at go-around
airspeeds, greater acceleration will result with
20 degrees of flaps than with takeoff position
of 30 degrees. Any flap setting lower than the
takeoff setting should be reduced to at least
the takeoff position immediately after de¬
cision to go around has been made.
5. Landing gear lever—UP. (P)
Note
® If an immediate obstacle must be cleared, do
not retract gear until obstacle is cleared. Re¬
traction of the gear creates considerable addi¬
tional drag. If the airplane is accelerating and
no immediate obstacle must be cleared, the
gear should be retracted.
# Landing gear should be up and locked and the
light in the control handle out before exceed¬
ing the structural limit airspeed,
6. Allow airplane to accelerate to 160 knots IAS
before attempting dim bout. If possible, stay close to
the runway to take advantage of "ground effect/ 1 (P)
A loss in airspeed of 5 to 10 knots should be anticipated
in leveling out. Should the airspeed be below 145 knots
before roundout, the airplane should be allowed to
touch down and accelerate on the runway. If a go-
around must be made after airplane touchdown, ac¬
celerate to a minimum of 135 knots IAS before liftoff
is attempted. If runway distance is available, attain
more than 135 knots before liftoff. After takeoff and
if conditions allow, take advantage of "ground effect* 1
by holding the airplane close to runway.
SINGLE-ENGINE TAKEOFF.
Single-engine takeoffs are not recommended for this
airplane. A single-engine takeoff chart (figure A-6) is
shown in the Appendix, but this chart is to be used
only for reference if an engine fails during takeoff.
SIMULATED SINGLE-ENGINE FLAMEOUT.
A single-engine flameout may be simulated by retard¬
ing the throttle (on simulated inoperative engine) to
IDLE, opening the throttle on the operating engine as
required, and placing the speed brake lever at 1/8 of
quadrant travel (opening speed brakes 15 degrees).
SIMULATED FORCED LANDING,
A two-engine flameout may be simulated for practicing
forced landings by retarding the throttles to 85% rpm
and opening the speed brake lever approximately 70
degrees (see figure 3-6).
3-9
Section III
T.O. 1F-89H-T
CHECK INSTRUMENTS FOR DESIRED
RANGES,
TURN ONTO FINAL AT 170 KNOTS JAS
AND STABILIZE AT 160 KNOTS |AS
i TAKEOFF FLAPS) OR 169 KNOTS END FLAPS)
DUMP TIP FUEL AND DROP
PYLON TANKS IF CARRIED.
ACTUATE SUPPLEMENTAL
HYDRAULIC PUMP.
RETARD THROTTLE TO IDLE ONLY
WHEN POSITIVE OF LANDING,
MAINTAIN HIGH ENGINE RPM
THROUGHOUT APPROACH-
DO NOT REDUCE AIRSPEED
BELOW ISO KNOTS UNTIL
LANDING IS ASSURED-
EXTEND DOWNWIND LEG TO ALLOW
LOWER THAN NORMAL FINAL APPROACH
ANGLE. THIS WILL PERMIT A H IGHER
RPM TO BE USED.
STABILIZE AIRSPEED AT 180 KNOTS.
TRIM—ADJUST AS AIRSPEED IS REDUCED.
H-II&DB
3-10
Figure 3-5,
TO. TF-89H-1
AIRSPEED^tSS KNOTS,
SPEED BRAKE LEVER-CLOSED,
landing GEAR lever—DOWN.
CHECK GEAR VISUALLY AND
WITH INDICATORS.
jsecfsosrc lui
NOTE:
Typical landing weight is bused on a typical area
intercept mission. Weight includes fuel for
20-mimtte loiter at sen level plus 5 percent total
fuel and full armament.
Increase landing speed 2 knots above speed cited ot
this landing chart for each additional WOO pounds
increase in gross aright.
note:
If go-around appears
necessarymake decision
us soon as possible.
WING FLAP LEVER—TAKEOFF,
IF FLAPS ARE AVAILABLE,
SPEED BRAKE LEVER — OPEN,
JF SPEED BRAKES AVAILABLE
ENTER TRAFFIC PATTERN AT
275 KNOTS IAS USING 91-93<?'o RPM.
APPLY WHEEL BRAKES INTERMITTENTLY
TO AVOID LOCKING WHEELS.
TURN EMERGENCY BRAKING
AIR ON IF NECESSARY.
SPEED BRAKE LEVER-AS REQUIRED.
IF SPEED BRAKES AVAILABLE,
SET NOSE WHEEL DOWN BEFQkt
REACHING T14 KNOTS IAS
(TAKEOFF FLAPS i OR
128 KNOTS (NO FLAPS L
WITH TAEL SLIGHTLY DOWN, TOUCH
MAIN WHEELS DOWN AT
*19 KNOTS (TAKEOFF FLAPS i OR
34 KNOTS (NO FLAPS).
uiRiwm;
-it higher gross weights, approach and touchdown
speeds must he increased. See landing speeds chart in
appendix for other weights and speeds. (Landing
speeds for TAKEOFl flaps are not shown in the
appendix. They are approximately 2 knots I AS higher
than for full flaps at all weights; nose wheel down
speeds are approximately l knot higher,)
lAo not extend flaps beyond TAKEOFF setting for a
single-engine landing, as the airplane will not
accelerate with full flaps and one engine operating
H-l 16(2)0
3-11
T.O. 1F-89H-1
DESCEND IN A STABILIZED SPIRAL
FORCED LANDING
hi MILE!
(appro:
7000 FEET”
155 KNOTS IAS
x
3000 FEET-
140 KNOTS IAS
NOTE: The above speeds apply to all gross weights of the airplane.
NOTE: All landings are to be made gear
down. If terrain is unknown or unsuitable for
a forced landing, eject in preference to
uttem pting a forced landing.
LOW KEY
11 FINAL APPROACH
SPEED-140 KNOTS IAS.
DROP PYLON TANKS IF CARRIED
BATTERY 12
SWITCH-OFF .
BEFORE TOUCHDOWN.
13
INERTIA REEL—-LOCKED
H-54E
WARN RADAR OBSERVER.
WARNING: • Battery life is extremely short when operating
the supplemental pump with no generator current available.
To avoid turning the pump on when lowering gear, turn
battery switch off and lower gear by emergency means
if all three generators are out
• If the battery is in poor condition, it may not support
supplemental pump operation. In this case, pressure
from windmilling engines must be depended upon for
final approach and flareout.
3-T2
Figure 3-6.
Changed 13 February 1959
T.O. 1F-S9H-1
Section 111
FIRE*
ENGINE FIRE DURING START,
If an engine overheat warning light comes on, close
both throttles and keep affected engine windmill ing
with ground test switch. If the light does not go out,
if an engine fire warning light comes on, or if there is
any other indication of fire, proceed as follows:
1. Engine fire selector switch for engine on fire—
Raise guard and actuate switch. (P)
2. Agent discharge switch—Actuate. (P)
3. Starter switch—STOP momentarily. (P)
4. Battery switch—'OFF. (P)
5* Generator switch(es)—OFF, (P)
6. Radar observer—Warn to abandon airplane, (P)
7, Abandon airplane as quickly as possible, {P—RO)
Do nor restart engine until cause of fire or
overheating has been determined and cor¬
rected. Never restart if agent discharge switch
has been actuated; this is a "one-shot” system,
and until the extinguishing agent bottle has
been replaced, there will be no further fire
protection available.
ENGINE FIRE DURING FLIGHT*
If an engine overheat or fire warning light comes on,
immediately retard the throttle on the affected engine.
If the light then goes out, continue flight with reduced
power and land as soon as possible. If either light stays
on, or if there is any other indication of fire, proceed
as follows:
1. Throttle—CLOSED (on inoperative engine), (P)
2. Engine fire selector switch {engine on fire)—
Raise guard and actuate. (P)
3. Agent discharge switch—Actuate. (P)
4. Radar observer—Alert, (P)
3. Oxygen diiuter lever—100% OXYGEN. (P—RO)
6. Oxygen regulator emergency lever—Actuate mo¬
mentarily, to clear oxygen mask. (P—RO)
Repeated or prolonged exposure to high con¬
centrations of bromochloromethane (CB) or
decomposition products should be avoided.
CB is a narcotic agent of moderate intensity
but of prolonged duration. It is considered
to be less toxic than carbon tetrachloride,
methylbromide, or the usual products of
combustion. In other words, it is safer to use
than previous fire extinguishing agents. How¬
ever, normal precautions should be taken in¬
cluding the use of oxygen when available.
7. Generator switch(es) (for inoperative engine)—
OFF. (P)
8. Do not attempt to restart engine. (P)
9. Land as soon as possible. (P)
The "one-shot” fire extinguishing system de¬
livers its entire charge when actuated and
must be recharged before it is used again,
FUSELAGE, WING, OR ELECTRICAL FIRE*
If fuselage, wing, or electrical fire occurs, perform the
following immediately:
L Radar equipment—Off. (F—ROj
2. All electrical equipment—Off. {P—RO)
3. Eject—If necessary. (P—RO)
SMOKE AND FUMES ELIMINATION.
1. Cabin air switch—RAM & DUMP. (P)
2. Oxygen diiuter lever—100% OXYGEN. (P—RO)
3. Oxygen regulator emergency lever—Actuate mo¬
mentarily, to clear oxygen mask. (P—RO)
EJECTION*
Escape from the airplane should be made with the ejec¬
tion seat. Follow the procedure shown in figure 3-7.
Ejection at airspeeds ranging from stall speed to 525
knots IAS results in relatively minor forces being ex¬
erted on the body. Ejection at airspeeds of 525 to 600
knots IAS exerts greater forces on the body, making
escape more hazardous. Above 600 knots, ejection is
extremely hazardous because of the excessive forces on
the body. Ejection at low altitude is facilitated by pull¬
ing the nose of the airplane up above the horizon in a
"zoom up” maneuver. Ejection seat velocity is small
compared to the velocity of the airplane so that ejec¬
tion accomplished when the airplane is flying horizon¬
tally results in the ejection seat following a nearly
horizontal path. A "zoom up” maneuver will result in
the ejection seat trajectory coming closer to the verti¬
cal, thus effecting an increase in altitude. This altitude
gain will increase the time available for separation
from the seat and deployment of the parachute. When
emergencies necessitate ejections, slow the airplane
3-13
Section III
T.O. 1F-89H-T
SIT ERECT WITH HEAD BACK, CHIN
TUCKED IN, BOTH ARMS ON ARMRESTS
AND FEET FIRMLY ON FOOTRESTS.
GRASP LOOPED
(RIGHT) HAND GRIP
AND PULL UPWARD.
GRASP LOOPED
(LEFT) HAND GRIP
AND PULL UPWARD
Minimum safe level flight attitude for ejection is 2000 feet
for manually operated safety belt and parachute; 1000 feet
for automatic safety belt and manual parachute; 1000 feet
with a manual belt opened before ejection and any type of
parachute; and 500 feet with an automatic belt and automatic
parachute (provided hey attached to the parachute timer
lanyard is inserted in the belt automatic release )*
Figure 3-7.
3-14
SQUEEZE FIRING
TRIGGER-RIGHT
HAND GRIP.
AFTER seat catapult fires-
SEAT BELTS AND SHOULDER HAR¬
NESS ARE UNCOUPLED AUTOMAT¬
ICALLY BY A DELAY INITIATOR.
WARNING
If time and conditions permit, the radar
observer rather than the pilot shall
jettison the canopy. This will assure
that the radar observer is in position
for ejection and will have no difficulty
in reaching the ejection seat controls
due to wind blast or li C” forces.
AFTER
EJECTION
(WITH SAFETY
BELT RELEASED)
KICK FREE OF SEAT.
WARNING
Keep hands and arms clear of canopy
control levers during canopy jettison.
As the canopy is jettisoned , the RO's
control lever will rotate rapidly to the
open position and the pilot’s control
lever will snap to the VP (open)
H-55(2)B
3-15
Section BID
T.O, 1F-89H-T
down as much as possible. Minimum safe ejection alti¬
tudes are 2000 feet with a manual belt and parachute,
1000 feet with an automatic belt and manual parachute,
1000 feet with a manual belt opened before ejection
and any type of parachute, and 500 feet with an auto¬
matic belt and automatic parachute (if the key attached
to the parachute timer lanyard is inserted in the belt
automatic release).
WARNING
Ejection should not be delayed when the air¬
craft is in a descending attitude and cannot be
leveled out. The chances of successful ejection
at low altitudes under this condition are
greatly reduced.
Minimum safe ejection altitudes for "one and zero”
and "two and zero" systems for various combinations
of equipment are listed below.
If the detachable lanyard has been installed
before the 1-second safety belt initiator, a
"two and zero' 1 system is temporarily pro¬
vided wherein higher minimum safe ejection
altitudes must be observed. (See following
table.)
For nonautomatic parachutes used with automatic
safety belts, lanyard. Part No, 57C620O, will be used.
The minimum safe escape altitudes specified for 1- or
2-second safety belt and 0-second parachute setting
apply when the lanyard is attached to the rip cord and
safety belt,
1 -Second 2-Second
Automatic Automatic
Lap Belt
Lap Belt
(Mil
(M4
Initiator)
Initiator)
2-Second Parachute
(F-lA Timer), B-4 or B-5
Pack, C-9 Canopy
350 Feet
55(?Feet
2-Second Parachute
(F-lA Timer), B-5 Pack,
C-ll Canopy
400 Feet
600 Feet
1-Second Parachute
(F-1B Timer), B-4 or B-5
Pack, C-9 Canopy
200 Feet
350 Feet
3-16
1-Second Parachute
(F-1B Timer), B-5 Pack,
C-ll Canopy
0-Second Parachute
(Lanyard to "D” Ring),
B -4 or B-5 Pack,
C-9 Canopy
0-Second Parachute
(Lanyard to "D” Ring),
B-4 or B-5 Pack,
C-ll Canopy
250 Feet 400 Feet
100 Feet 200 Feet
150 Feet 250 Fee
BEFORE EJECTION.
1. Reduce airspeed as much as possible and, if below
2000 feet, pull the nose up above the horizon to reduce
airspeed ("zoom up” maneuver). (P)
2. Pull handle on bailout bottle if altitude necessi¬
tates, (P—RO)
3. Cabin air switch—RAM & DUMP. (P)
4. Loose equipment—Stow. (P—RO)
5. Automatic-opening parachute—Check. (P—RO)
Make sure the key is attached to the automatic-
opening safety belt and the lanyard is free,
6. Canopy—Jettison. (P—RO)
Keep hands and arms clear of canopy lock
levers during canopy jettison. As the canopy
is jettisoned, the radar observer's lock lever
will rotate rapidly to the open position, and
the pilot's lock lever will snap to the up
(open) position.
EJECTION PROCEDURE.
1. Left handgrip—Grasp and pull upward, (P—RO)
2. Right handgrip—Grasp and pull upward.
(P—RO)
3. Firing trigger (on right handgrip)—Squeeze.
(P—RO)
4. After ejection (with safety belt released)—Kick
free of seat. (P—RO)
FAILURE OF SEAT TO EJECT.
If the ejection seat fails to operate, the following pro¬
cedure may be used for leaving the airplane:
1, Reduce speed, (P)
2, Oxygen hose, radio equipment, and "G” suit lines
—Disconnect. (P—RO)
3. Safety belt—Release. (P—RO)
4. Bailout—If possible roll airplane on its back and
push dear. If it is not possible to roll the airplane over,
bail out the side of cockpit toward the aft trailing edge
of the wing. (P—RO)
T.O. IF-S9H-1
Section HI
FAILURE OF CANOPY TO JETTISON,
L Canopy jettison "T” handle—PulL (P)
2. Canopy lock lever—Raise (if step 1 is ineffective),
(P—RO)
3. Canopy switch—Move to OPEN (if step 1 is inef¬
fective). (P —RO)
Note
As the canopy moves aft from the windshield
frame, the airstream will blow it from the
fuselage,
4. Continue with normal ejection procedure and
eject through canopy (if steps 1,2, and 3 are ineffec¬
tive). (P—RO)
AFTER EJECTION.
L After safety belt releases automatically, kick away
from seat. (P—RO)
Note
If automatic release fails, release safety belt
manually.
2, Conventional parachute—Delay opening to allow
body to decelerate so that opening shock will be re¬
duced. (P—RO)
Note
If ejection is at high altitude, free fall to nor¬
mal breathing altitude will reduce the dan¬
gers of hypoxia and cold.
• With a manual opening safety belt, open the
belt before ejection under the following con¬
ditions: below 2000 feet if position in seat
can be maintained, in a dive, of at high air¬
speeds up to 5000 feet. Opening the belt
before ejection will facilitate pulling the D-
ring, and parachute opening after ejection.
At altitudes higher than 5000 feet do not open
manually operated safety belt before ejection,
especially at high airspeeds,
® With an automatic opening safety belt do not
open the belt before ejection at any altitude.
The automatic opening feature will give you
the maximum safety factor under all condi¬
tions.
In all ejections below 14,000 feet, manually
pull the parachute D-ring immediately fol¬
lowing separation from the ejection seat. This
applies regardless of parachute type, manual
or automatic.
LANDING EMERGENCIES,
LANDING WITH LATERAL UNBALANCE AND CRIT¬
ICAL AFT CG.
When fuel remains in tip tanks, with no rockets or
ballast in the nose of the rocket-missile pods, and with
the right main tank nearly empty, a condition of crit¬
ical aft eg may exist. If, in connection with this, a
condition of lateral unbalance is suspected or known
to exist, do not attempt rolling maneuvers, and use
minimum bank turns only, in preparation for landing.
Use the following procedure for landing with lateral
unbalance and critical aft eg:
1. Tip tank fuel—Dump and burn off fuel unbal¬
ance as much as possible, (P)
2, Speed brakes—Use approximately 15 degrees for
roll control. There will be no drag penalty and aileron
effectiveness will be increased approximately 30 per¬
cent, (P)
3, Make straight-in approach. (P)
4. If time and fuel are available, test the airplane in
the intended landing configuration at a minimum alti¬
tude of 12,000 feet to determine how slowly the air¬
plane may be flown in a wings level attitude using a
maximum of 1/2 aileron throw. Touchdown should be
made at this speed. If adequate runway exists, plan to
land without flaps for better aileron control. (P)
LANDING WITH ONE TIP TANK CONTAINING
FUEL.
If fuel from one tip tank cannot be dumped using nor¬
mal and emergency fuel dumping procedures, the fol¬
lowing procedure for landing with an asymmetrical
tip fuel condition will be used:
1. Maintain sufficient airspeed to fly the airplane in
wing-level attitude using maximum of one-half aileron
throw. (P)
To provide a margin of safety, aileron deflec¬
tion should be limited to approximately one-
half aileron throw to maintain wing-level
flight. For minimum recommended approach
a irspeeds with asym metr i cal tip fuel co nd i-
tion, see Asymmetrical Tip Fuel Condition
VS Airspeed chart (figure 3-8).
2. Trim—-Use trim switch to streamline ailerons
with wings as an initial setting to indicate the amount
of aileron control remaining. (P)
3-17
3-18
INDICATED AIRSPEED-KNOTS
T.O. 1F-89H-1
Section III
THIS PAGE INTENTIONALLY LEFT BLANK
Changed 13 February 1959
3-18 A
Section III
T.O. 1F-89H-1
# Aileron trim switch should be used only to
streamline ailerons with wings as an initial
setting,
& Avoid turning maneuvers as much as possible,
holding roll rate to absolute minimum,
® Bank angle is limited to 30 degrees maximum
in either direction; however, where possible,
turns should be made to the side with the
least fuel.
3. Pylon tanks—Jettison (if carried), (P)
4, Approach—Make straight in, (P)
LANDING WITH FLAPS AND SPEED BRAKES
RETRACTED.
If wing flaps and speed brakes are unavailable for land¬
ing, higher touchdown airspeeds must be used to com¬
pensate for the lack of extra lift normally supplied by
the flaps. If both engines are operative, use speed
brakes (if available) and maintain a minimum of 85%
rpm throughout the final approach for rapid accelera¬
tion if a go-around is necessary. Lengthen the down¬
wind leg to provide for a flat final approach and main¬
tain engine rpm at as high a setting as possible. The
airplane should be flown strictly by the airspeed indi¬
cator throughout the final approach and touchdown
Recommended final approach speeds are as follows:
Gross Weight—Lb
30,000
34,000
38,700
Approach IAS—Knots
160
169
179
Note
The preceding speeds are approximately 5
knots above the stall speed encountered under
average gust conditions.
Touch the main wheels down at the following appli¬
cable airspeed:
Grass Weight — Lb
30,000
34,000
38,700
Landing IAS—Knots
125
133
142
Set the nose wheel down before the following appli¬
cable airspeed is reached:
8-18B
Changed 13 February 1959
TO. 1F-89H-1
Section 111
Gross W eight — Lb Nose W heel Do ivnlAS—Knots
30,000 119
34,000 127
38,700 140
Anticipate a landing roll 25 to 35 percent longer than
normal for a dry hard-surfaced runway.
RUNWAY OVERRUN BARRIER OPERATION
(SOME AIRPLANES).
On some airplanes, a runway overrun barrier engage¬
ment modification is provided to prevent injury to
crewmembers and damage to equipment. This modi¬
fication prevents airplanes from overrunning runways
if pilot should be unable to stop the airplane during
landings or unsuccessful takeoffs. Airplanes so modi¬
fied are equipped with a barrier triggering probe
located just forward of the nose wheel gear well and
a barrier guide and hook located just forward of the
main landing gear. When the airplane overruns the
end of the runway, the probe triggers the barrier,
actuating barrier to an upright position. The guide
then guides the barrier to the engaging hook which
arrests the forward momentum of the airplane. A test
program showed that the airplane could be successfully
arrested at speeds ranging from 29 to 83 knots with
all wheels making firm contact with the ground. This
system is operable when the runway is equipped with
Type MA-1A runway overrun barriers.
RUNNING OFF RUNWAY ON LANDING.
During the landing roll, if the airplane leaves the run¬
way due to failure of brakes, failure of arresting gear
to engage the MA-1A runway overrun barrier on air¬
planes so equipped guide the airplane towards the
smoothest terrain, if possible. The landing gear may
he either raised or left extended. If desired, the land¬
ing gear can be raised by depressing the emergency
override lever and simultaneously moving the landing
gear lever to UP.
FORCED LANDING.
If it is necessary to make a forced landing, accomplish
as much of the procedure shown below and on figure
3-6 as possible. Land with the gear down regardless of
the terrain, as statistics prove that less personal injury
and damage to equipment are likely to result from a
gear-down landing. Two-engine flameout landings
should he considered only under most favorable or
extenuating circumstances.
Note
If landing is to be made on a runway, leave
landing gear lever in the DOWN position
if the runway is equipped with Type MA-1A
runway overrun barrier and the aircraft is
modified to contain the necessary arresting
gear equipment.
Changed 13 February 1959
# Battery life is extremely short when operating
the supplemental pump with no generator
current available. To avoid turning the pump
on when lowering gear, turn battery switch
off, if all three generators are out, and lower
gear by emergency means,
$ If battery is in poor condition, it may not sup¬
port supplemental pump operation. In this
case, pressure from wind milling engines must
be depended upon for final approach and
landing. Rapid movement of flight controls
must be avoided.
® Do not raise the helmet visor prior to landing
emergencies. Retaining the helmet visor in
the lowered position will afford protection
from impact injury, dislodged objects in the
cockpit, flame and smoke, and from wind¬
blown objects if the canopy is jettisoned. The
helmet visor is a protection device that should
be worn in the lowered position in all land¬
ing emergencies.
Note
A two-engine flameout may be simulated for
practicing forced landings by retarding the
throttles to 85% rpm and opening the speed
brake lever approximately 70 degrees (see
figure 3-6).
1. Radar observer—Warn of impending forced land¬
ing. (P)
2. Pylon tanks—Jettison. (P)
3. Tip tank fuel—Dump. (P)
4. Throttles—CLOSED if power failure is complete;
otherwise, use available power. (P)
5. Landing gear lever—DOWN at 10,000 feet. (P)
Note
If landing is to be made on a runway, leave
landing gear lever in the DOWN position
if the runway is equipped with Type MA-IA
runway overrun barrier and the aircraft is
modified to contain the necessary arresting
gear equipment.
Battery life is extremely short when operat¬
ing the supplemental pump with no genera¬
tor current available and may not last long
enough to permit a two-engine flameout land¬
ing, To avoid turning the pump on when low¬
ering gear, turn battery switch off and lower
gear by emergency means.
3-19
Section ill
T.O, 1F-89H-1
6, Wing flap lever—-TAKE OFF. (P)
7, Parachute—Unbuckle, and safety belt—Tighten,
(P—RO)
8, Engine fire selector switches—Actuate, (P)
9, Generator switches—OFF. (P)
10. Canopy—Jettison with "T” handle, (P)
11- Final approach airspeed—140 knots IAS. (P)
Airspeeds on final approach and flareout can¬
not be depended upon to windmill the engines
at sufficient rpm to maintain hydraulic pres¬
sure for flight operation,
12. Battery switch—OFF. (P)
13. Inertia reel—LOCKED, just before touchdown,
(P—RO)
I; CAUTION ;[
When the shoulder harness inertia reel is
locked, either by use of the inertia reel lock
lever or by raising the left armrest, the pilot
is prevented from bending forward; there¬
fore, all controls not readily accessible should
he positioned before the inertia reel is locked.
LANDING WITH GEAR PARTIALLY EXTENDED.
If the landing gear fails to extend completely after
both the normal and the emergency procedures have
been used, leave as many wheels down as will extend,
jettison canopy with "T” handle, and proceed with
forced landing. Less damage will result with this pro¬
cedure than with a gear-up landing.
LANDING WITH FLAT TIRE.
Because of the comparatively large diameter wheels
and small width of tires, directional control of the air¬
plane is easily maintained with rudder and wheel
brakes if a main gear tire blows out on landing. If the
airplane is landed with one nose wheel tire flat, there
will be a slight veering. If a landing is made with both
nose wheel tires flat, sufficient up-elevator should be
applied to take weight off the nose wheel, and use of
wheel brakes should be minimized.
EMERGENCY ENTRANCE*
If it is necessary to gain entrance to the cockpit in an
emergency, first attempt to open the canopy by using
the external lock lever and canopy switch, both located
behind an access door on the left side of the fuselage
above the wing leading edge. If the canopy switch fails
to open the canopy after it is unlocked, attempt to open
it manually using the external handgrips in the aft
structure of the canopy. If this fails, slow-jettison the
canopy by use of the external emergency canopy release
handle, located flush with the fuselage skin just below
the access door for the external canopy switch. Pushing
the button in the center of the handle will release it.
The handle must be pulled out (with a force of about
45 pounds) approximately 5 inches and held (from 10
to 20 seconds) until the canopy is raised above the cock¬
pit rails. The canopy can then be lifted or pushed from
the airplane. If all of the foregoing procedures fail,
then, as a last resort, chop a hole in the canopy with an
ax, using extreme caution not to injure crewmembers
inside the cockpit. On airplanes modified in accord¬
ance with T.O. IF-89-659, new rescue markings will
be visible on either side of the cockpit. See figure 1-40.
EMERGENCY EXIT ON GROUND*
If canopy cannot be opened by the normal procedure
and immediate exit is necessary, in an emergency the
radar observer can slow-jettison the canopy by using
the emergency hydraulic pump handle to put pressure
against the cable attached to the canopy external jetti¬
son lever and the canopy jettison initiator. When pres¬
sure is applied to the cable between the control lever
and the initiator, the initiator is actuated and the can¬
opy is slow-jettisoned. The cable is located in the for¬
ward left side of the radar observer's cockpit. If the
canopy cannot be slow-jettisoned, fast-jettison the can-
opy by pulling the canopy jettison handle (on the
pilot's right vertical console) or by raising the right
armrest of either ejection seat. On airplanes modified
in accordance with T.O. 1F-89-586, both the pilot's
and radar observer's cockpits are equipped with an
internal canopy slow fire jettison "T” handle. This
enables either the pilot or the radar observer to slow-
jettison the canopy by pulling the "T” handle.
General direction of canopy movement when
fast-jettisoned is straight up. Lack of atrstream
may cause it to fall back into the cockpit,
DITCHING.
This airplane should never be ditched if there is suf¬
ficient altitude for safe ejection. Ditching is not
recommended because it is assumed that the engine air
intake ducts will cause the airplane to dive violently
when it hits water. However, if altitude is insufficient
for ejection, warn radar observer, then proceed as
follows:
L Tip tank fuel—Dump, (P)
3-20
T.O. 1F-89H-T
Section ill
Empty tip tanks do not contact the water
until the airplane comes to rest r where they
afford additional buoyancy. If the tip tanks
contain fuel on ditching, they may plane
through the water and create serious decelera¬
tion loads. Pylon tanks should be jettisoned
whether or not they contain fuel.
2. IFF master control knob—EMERGENCY. (P)
3. External stores—Jettison, (P)
4. Canopy—Jettison with "T” handle. (P)
5. Landing gear lever—UP. (P)
6* Safety belt—Tighten. (P—RO)
7. Oxygen diluter lever—100%. (P—RO)
8. Nose wheel steering button—Depress. (P)
9. Wing flap lever—TAKE OFF. (P)
10, Throttles—CLOSED* (P)
XL Engine fire selector switches for both engines—
Raise guards and actuate. (P)
12, Select a heading parallel to the wave crest if pos¬
sible. Try to touch down along wave crest or just after
crest passes, (P)
13* Make norma! approach. (P)
14, Flare out to landing attitude, keeping the nose
high. (P)
15, Generator switches—OFF, (P)
16. Battery switch—OFF, just before contact. (P)
17. Inertia reel—LOCKED. (P—RO)
♦ After battery switch is turned OFF, supple¬
mental pump will not operate.
Do not attempt to ditch in a near level atti¬
tude. It is assumed that the airplane will dive
violently when the intake ducts hit the water.
OIL SYSTEM FAILURE.
If loss of oil is experienced, the airplane need not be
abandoned immediately as a gas turbine engine will
not fail immediately after loss of oil. An airplane gas
turbine engine depends upon oil to cool the roller and
hall bearings, so in the event of oil loss, reduce power
to keep temperatures at a minimum. A J35 engine has
operated for 27 minutes without oil before experienc¬
ing destructive engine failure. In most instances, ulti¬
mate failure of the engine will not occur within 10
minutes after loss of oil and will be characterized by a
steadily increasing vibration. At this time engine shut¬
down should be made to prevent a destructive engine
failure that would jeopardize a successful ejection or
power off control of the airplane in a landing attempt.
In most cases the airplane has remained controllable
during its descent. When oil loss is experienced, the
following procedure should be performed immediately:
1. Tip tank fuel—Dump. (P)
2. "G” forces—Minimize. (P)
3- Power setting (affected engine)—Minimum. (P)
4. Land at nearest airbase. (P)
FUEL VENT SYSTEM MALFUNCTION.
Under certain conditions of fuel vent or transfer system
malfunction, fuel may be lost overboard through a
main tank vent. If fuel overboarding occurs, use the
corrective procedures described in the following para¬
graphs.
FUEL OVERBOARDING DURING CLIMB OR DIVE.
Overboarding of fuel during a climb or dive indicates
mechanical failure of the left main tank dive valve in
the open position. When this condition occurs, the air¬
plane should be leveled immediately. If the fuel over¬
boarding stops, malfunction of the dive valve is con¬
firmed. If another climb or dive is made, and fuel
overboarding starts again the airplane should be lev¬
eled immediately, and the inboard and outboard wing
tank pump circuit breakers pulled out (deenergized)
until the main tank low-level warning light illumi¬
nates. Both wing tank pump circuit breakers should
then be pushed in (energized). If fuel overboarding
continues, repeat the operation as necessary. If this
procedure fails to stop fuel overboarding, wing tanks
should be selected and the airplane landed as soon as
possible.
If fuel overboarding cannot be stopped, it is
recommended that a no-flap landing be made
because overboarding fuel could be drawn
into the flap wells and drain into the hot en¬
gine bay, resulting in a fire or explosion.
Note
Since the vents for both the left and right
main tanks are located adjacent to each other,
the tank that is overboarding fuel cannot be
determined.
3-21
Section III
TO. 1F-S9H4
FUEL LEVEL CONTROL SHUTOFF VALVE
MALFUNCTION.
If water freezes in a main tank fuel level control shut¬
off valve during normal fuel sequencing (ALL TANKS
selection), the valve will not close fully when the main
tank is full. As a result, fuel is forced overboard
through the main tank vent line. (This condition will
also exist if foreign matter other than ice accumulates
in the main tank shutoff valve.) To correct this condi¬
tion, the fuel selector switch for the left fuel system,
then the right fuel system should be turned from ALL
TANKS to WING TANKS and back to ALL TANKS.
The resultant surges of fuel from the wing tanks may
free the valve in the malfunctioning main tank. If the
valve remains stuck, a wing tank selection should be
made and the airplane landed as soon as possible.
If fuel overboarding cannot be stopped, it is
recommended that a no-flap landing be made
because overboarding fuel could be drawn
into the flap wells and drain into the hot en¬
gine bay, resulting in a fire or explosion.
FUEL SYSTEM EMERGENCY OPERATION.
FUEL SYSTEM OPERATION FOLLOWING COMPLETE
ELECTRICAL FAILURE.
Without electrical power, fuel is only available by
gravity feed from the tank selected directly to the en¬
gine. Fuel will not transfer from the wing tank to the
main tank without a booster pump. With battery
power available for a limited time during complete
electrical failure, the best fuel selection to ensure
gravity feeding can be obtained using the following
procedures:
L Right fuel selector switch—ALL TANKS. Engine
power settings when utilizing gravity feed should not
exceed those required for maximum endurance as listed
in the appendix. The maximum altitude at which the
right main tank can maintain satisfactory engine op¬
eration on gravity feed is approximately 25,000 feet.
If it becomes necessary to extend the flight
beyond the limits of the fuel available in the
right main (nose tank), the battery switch
should be placed at ON and right fuel selector
switch placed at WING TANKS. Fly the air¬
plane in a level flight attitude.
2. Left fuel selector switch—WING TANKS (left
wing tanks and main tanks will feed fuel to the en¬
gines simultaneously). The left wing should be elevated
to ensure maximum fuel flow from the wing tanks. The
maximum altitude at which the left main tank and
wing tanks can maintain satisfactory engine operation
on gravity feed is approximately 18,000 feet. (P)
;! CAUTION
Attempts to gravity feed above the foregoing
maximum gravity feed altitudes may cause
damage to the engine-driven fuel pump which
will result in fuel leakage.
3. Crossfeed switch—OFF. (P)
Crossfeed switch must be placed at CLOSED
during gravity feed operation. This will pre¬
vent flameout of both engines due to one main
tank emptying prior to the other.
MAIN TANK BOOSTER PUMP FAILURE.
If one of the two main tank booster pumps fails, the
remaining pump will continue to supply fuel to the
engine and afterburner above 30,000 feet. Depending
upon atmospheric conditions, one pump may or may
not support afterburning below 30,000 feet.
WING TANK BOOSTER PUMP FAILURE.
Wing tank booster pump failure will be evidenced by
wing heaviness or main tank low-level warning light
glowing.
Note
Under some conditions of speed, altitude, and
temperature, afterburning may not be avail¬
able on WING TANKS selection below
10,000 feet; therefore WING TANKS selec¬
tion is not recommended for takeoff. A
WING TANKS selection will support after¬
burner operation above 10,000-foot altitude;
however a wing tank booster pump failure
will cause the related afterburner to shut
down without warning when operating on
WING TANKS selection below approxi¬
mately 30,000-foot altitude.
If pump failure is caused by an electrical overload
condition, the related circuit breaker on the fuel con¬
trol panel (figure 148} will pop OUT, indicating
which pump has failed. If no circuit breakers are OUT
and a pump is believed to be inoperative as evidenced
by wing heaviness or low-level warning light, place
3-22
T.O. 1F-89H-T
Section III
the fuel quantity selector switch for the system with
the heavy wing alternately at INBD and OUTBD
while observing the related fuel quantity gage; a large
quantity of fuel existing in one wing tank after the
other wing tank in the same system is empty or nearly
empty, will indicate a pump failure in the tank with
the most fuel. To use fuel from a wing tank with an
inoperative pump, see Gravity Feed, this section.
GRAVITY FEED.
Each fuel system will gravity feed the related engine
with sufficient fuel to maintain military power up to
an approximate 6000-foot altitude. When fuel is being
gravity fed from either or both wing tanks, the air¬
plane must be flown at a 10-degree uncoordinated
wing-high attitude for the system that is gravity feed¬
ing, so that the maximum amount of fuel can be
drawn from the wing tanks. To gravity feed fuel
from either wing tank, the booster pump in the other
wing tank in the same system must be shut down by
pulling the related circuit breaker. After fuel supply
is exhausted from the wing tank during gravity feed¬
ing, press the circuit breaker IN.
Note
When gravity feeding from the right wing
tanks with the fuel selector at WING
TANKS, the right main tank is isolated from
the system. Caution must be exercised to avoid
flameout after the wing tanks become empty.
Use the fuel quantity gage system to antici¬
pate the wing tanks becoming empty; place
the fuel selector at ALL TANKS before com¬
pletely emptying the wing tanks.
ONE TIP TANK NOT FEEDING.
When fuel is not feeding from one of the tip tanks and
an asymmetrical fuel condition is indicated, use the
following procedure;
1. Fuel selector switch—ALL TANKS. (P)
2. Pylon tanks—Jettison {if carried). (P)
Nofe
If fuel is not being fed from a tip tank due to
lack of pressurization, the pylon tank on the
same side, in all probability, will not feed. Re¬
taining the pylon tanks will aggravate the
asymmetrical fuel condition.
3. If more than one-half aileron throw is required
to maintain a wing-level attitude, increase airspeed as
required to maintain level flight. Aileron trim should
be used only to streamline aileron with wings as an
initial setting to indicate the amount of aileron control
remaining. (P)
• To provide a margin of safety, aileron deflec¬
tion should be limited to approximately one-
half aileron throw to maintain wing-level
flight. For minimum recommended approach
airspeeds with asymmetrical tip fuel condi¬
tion, see Asymmetrical Tip Fuel Condition
VS Airspeed Chart (figure 3-8).
• Avoid turning maneuvers as much as possible,
holding roil rate to absolute minimum,
• Bank angle is limited to 30 degrees maximum
in either direction; however, where possible,
turns should be made to the side with the least
fuel,
4. Tip tank fuel—Dump by normal means if pos¬
sible. If normal fuel dumping is unsuccessful, place
the airplane in a moderate speed climb and dump the
fuel by gravity. This is done by pressing the tip tank
button which energizes both tip tank dump valves
and holds them open for approximately 75 seconds by
means of a time-delay relay. Several fuel dump cycles
will be required to empty the tip tanks. Approximately
5 minutes are required to complete dumping tip tank
fuel by gravity. (P)
ENGINE-DRIVEN FUEL PUMP FAILURE.
The engine-driven fuel pump has two individual
pumping elements. If either pumping element fails,
the remaining pumping element will continue to sup¬
ply adequate fuel pressure to operate the engine at
military power at any altitude. Failure of a pumping
element in an engine-driven fuel pump will not affect
afterburner operation. Fuel pressure for afterburning is
supplied by the booster pumps and afterburner tur¬
bine-driven pump. Warning lights located on the
pilot's left console (figure 1-9) will warn of a pump
element failure during ground operation. During
flight the warning lights are automatically disarmed
by a switch on the left main landing gear.
DAMAGED TANKS.
If tanks are damaged or a severe leak is suspected, take
corrective action as described in the following para¬
graphs.
Damaged Main Tank,
1. Main tank fuel—Use first. (P)
2. Fuel selector switch (for damaged fuel system)—
ALL TANKS. (P)
3. Wing tank booster pump circuit breakers (on side
where damage exists)—Pull. (P)
4. Wing tank booster pump circuit breakers (after
main tank fuel is used)—IN. (P)
5. Fuel selector switch (for damaged fuel system)—
WING TANKS. (P)
3-23
Section lit
TO. 1F-89H-T
Escaping fuel from a damaged right main
tank may enter the air intake duct of an
engine and cause the engine to explode.
Damaged Tip or Pylon Tank.
If a tip tank or pylon tank is damaged and pressure is
lost, fuel from the tip and pylon tanks will not be
usable. Jettison the pylon tanks and press the tip
tank fuel dump button. Two or more dumping cycles
may be required to dump fuel from unpressurized tip
tanks. Any fuel remaining in the forward section of
an unpressurized tip tank will not be dumped.
Damaged Wing Tanks.
If either wing tank is damaged and its booster pump
is operative, use available fuel in the damaged tank
first by pulling OUT the booster pump circuit breaker
of the undamaged wing tank in the same system. If
damage is such that escaping fuel causes an obvious
fire hazard, shut down the engine on the damaged
side.
AFT CENTER-OF-GRAVITY FUEL MOVEMENT
When fuel is in the tip tanks, weight of the right
main tank fuel is essential to keep the airplane's center-
of-gravity within allowable limits. If right main
tank fuel is lowered 50 gallons (325 pounds) from full
before tip tanks empty, an aft eg warning light on
the pilots instrument panel will come on to warn of
an aft eg condition caused by insufficient fuel being
transferred into the right main tank. If the aft eg
warning light comes on, the pilot must reduce air¬
speed to Mach 0,65 or below and reduce power on
the right engine. If the aft eg warning light stays on
and the right main tank fuel level continues to
drop as noted on the fuel quantity gages, immediately
place the right fuel selector at WING TANKS. When
all fuel is used from the tip tanks, as indicated by the
fuel quantity gage system, use remaining fuel in wing
and main tanks.
Nate
© The aft eg warning light is disarmed by a
float switch in the right tip tank when the
tank becomes empty. This float switch is only
to prevent the warning light from burning
after all tip tank fuel is expended. To deter¬
mine if any fuel remains in the tip tanks, use
the fuel quantity gage system.
# Sustained low-altitude operation at maximum
power can cause the rate of fuel consumption
from the main tanks to exceed the rate of re¬
plenishment from the wing tanks. If the aft
eg warning light comes on under these condi¬
tions, reduce power on the right engine or
increase altitude.
ELECTRICAL SYSTEM EMERGENCY
OPERATION.
See figure 1-21 for equipment rendered inoperative
because of failure of the 28-volt d-c system, the a-c
alternator system, or the a< inverter systems. In case
of complete electrical failure, do not abandon airplane
as control of the airplane can be maintained. Figure
3-9 covers 28-volt d-c generator malfunction for which
corrective action may be taken by the pilot.
GENERATOR OVERVOLTAGE.
If the voltage of a generator becomes excessive, an
overvoltage relay will cut the generator out of the
circuit and the generator warning light will come on.
To return the generator to the circuit proceed as
follows:
1. Generator switch—RESET momentarily, then re¬
turn to ON. (P)
If the generator warning light goes out and re¬
mains out, overvoltage was temporary.
2. Generator switch—OFF if generator warning
light remains on. (P)
3. Voltage regulator rheostat—Turn toward DEC to
reduce voltage, (P)
4. Generator switch—RESET momentarily, then re¬
turn to OFF. (P)
5. Voltage selector switch—Turn to affected genera¬
tor. (P)
6. Voltage regulator rheostat—Adjust until voltage
is slightly above the voltage of the other generators. (P)
7. Generator switch—ON. (P)
GENERATOR FAILURE.
If any one of the three 28-volt d-c generators fails,
the remaining two can carry the entire load. If any two
generators fail, the secondary bus is automatically dis¬
connected from the electrical system. The remaining
generator supplies power to the primary bus only. See
figure 1-21 for power distribution. If all three 28-volt
d-c generators fail, the following procedure is recorm
mended:
1. Battery switch—OFF, (P)
2. Battery switch—ON and OFF as required to con¬
trol the fuel system. (P)
3* All electrical switches not essential to emergency
flight—OFF, (P—RO)
4. Tip tank fuel—Dump (if required). Place battery
switch in the ON position, push tip tank dump button
and hold for 5 seconds, then place the battery switch in
the OFF position. (P)
The tip tank dump valve will remain open until
the battery switch is again placed in the ON posi¬
tion.
See figure 3-9 for procedure in case of 28-volt d-c mal¬
function.
3-24
t-VOlT D-C GEHERATOR
TO. 1F-S9H-1
Figure 3-9,
3-25
Section lit
T.O. 1F-B9H-T
ALTERNATOR FAILURE.
If the 115/200-volt alternator fails, all the components
powered by it will be inoperative except the IFF and
the windshield defog heat* These will be switched to
the single-phase inverter system and will remain in
operation. (See figure 1-21,)
switch at the other position. If the light comes on
again, indicating failure of both inverters, place the
three-phase inverter switch at OFF* See figure 1-21 for
equipment powered by the inverter systems*
INSTRUMENT FAILURE.
INVERTER FAILURE.
If the main single-phase inverter fails during normal
operation, as indicated by the essential bus warning
light coming on, place the single-phase inverter
switch at EMERGENCY; the spare inverter will then
assume the load of the essential bus and the secondary
bus will not be energized* If the essential bus warning
light comes on again, after emergency operation has
been selected, failure of the spare inverter is indicated
and the single-phase inverter switch should be placed
at OFF.
® If both single-phase inverters fail below 10,000
feet, or if the afterburner ac control circuit
breaker pops out, the afterburner and after¬
burner control circuits will be inoperative*
When this occurs, the throttle-actuated eye¬
lid switches will cause the eyelids to open
(without regard to afterburner operation)
when the throttles are advanced to OPEN,
resulting in very low tailpipe temperatures
and extreme loss of thrust. If both inverters
fail below 10,000 feet while in afterburning,
afterburner operation will be unaffected.
However, if the afterburners are shut down
by depressing the throttle fingerlifts, the eye¬
lids will remain open* The eyelids must be
closed by moving the afterburner control cir¬
cuit breakers to OFF or by retarding throttles
to approximately 90% rpm position* Eyelid
closure will be apparent by an immediate in¬
crease in thrust and a return to normal tail¬
pipe temperature* Only military power will
be available for the duration of the flight*
® If both single-phase inverters fail while in
afterburning above 10,000 feet, afterburning
will be unaffected because the holding relay
in the afterburner control box keeps the eye¬
lids open; however, once afterburning is shut
off, it cannot be reinitiated* If both single¬
phase inverters fail above 10,000 feet, after¬
burning cannot be initiated and eyelids will
remain in closed position, because the altitude
switch breaks dc operating circuit, allowing
the fail-safe eyelid control valve to keep the
eyelids closed*
If the selected (main or spare) three-phase inverter
fails, as indicated by the three-phase inverter warn¬
ing light coming on, place the three-phase inverter
Engine Instruments.
ENGINE
INSTRUMENT FAILURE
If both single-phase inverters fail, all engine instru¬
ments will become inoperative except the tachometers
and exhaust gas temperature gages, both self-gen¬
erating instruments* The pointers of the oil pressure
gages, fuel pressure gages, fuel quantity gages, fuel
flowmeter indicator, brake accumulator pressure gage,
and left and right hydraulic pressure gages, all powered
by the single-phase inverter system, will remain at the
last setting indicated before inverter failure unless
moved by vibration or shock*
Flight Instruments.
rm fiiGNTmm
INSTRUMENT FAIL URE
28-volt d-c system falls:
If 28-volt d-c system or all three-
3-26
T,<X 1F-89H-1
Section 111
If all electrical systems fail, the following instruments
will remain in operation: vertical velocity indicator,
airspeed indicator, standby magnetic compass, and
altimeter. The vertical velocity indicator, altimeter, and
airspeed indicator will operate as long as the inlets on
the pitot tube and static ports are not iced over. The
turn and slip indicator depends on 28-volt d-c power
for operation. If both three-phase inverters fail, the at¬
titude indicator will tumble. The gyrosyn compass sys¬
tem and flight computer are powered through a phase
converter by the single-phase essential bus. If both
single-phase inverters fail, then the gyrosyn compass
system and flight computer will receive power from the
three-phase inverter system. See figure 1-21 for equip¬
ment powered by the inverter systems. If the 28-volt d-c
system fails, the free air temperature gage needle will
fall against the stop and all instruments depending
upon power from either the single-phase or three-phase
inverter system will be inoperative,
HYDRAULIC SYSTEM EMERGENCY
OPERATION,
If the right hydraulic system fails, all hydraulically
operated units will operate by pressure from the left
hydraulic system; however, flight control operation
will be limited in degree and rate of surface movement.
To increase the degree and rate of control surface
movement during operation of other hydraulic units,
depress the nose wheel steering button to start the left
hydraulic system supplemental pump.
With right hydraulic system pressure un¬
available, do not operate speed brakes unless
the left engine rpm is at least 85% or the sup¬
plemental hydraulic pump is operating. At
lower rpm, the demand on the left hydraulic
system by speed brake operation results in
limited aileron control unless supplemental
pump pressure is available.
If the left hydraulic system fails through loss of fluid,
the flight control system will operate on the right
system pressure, but the degree and rate of surface
movement will be limited. Speed brakes will be inop¬
erative, The landing gear and wheel brakes (if accu¬
mulator pressure is not available) must be operated by
emergency procedures. If the failure is caused by
engine-driven pump failure only, system pressure can
be maintained with the supplemental pump, and all
units normally operated by the left system will be
available, although rate of response may be somewhat
less than normal. Operation of speed brakes, flaps, and
gear in rapid sequence should be avoided. Use of nose
wheel steering should be held to a minimum because I
of the high volume of fluid required. If partial failure I
occurs through the failure of an engine but the engine I
is still windmilling, pressure can be expected to vary I
between 700 and 2000 psi. Care should be taken not I
to allow the pressure to bleed below approximately I
600 psi. This allows a slight margin above the purge I
valve setting of 300 psi. When this valve opens, pump |
flow is routed to the return line with the resultant loss
of the system. The only means for closing the valve
would be to increase engine rpm to about 38% or
energize the supplemental pump. Engine windmill
speeds to be expected are approximately 16%, 12%, and
9% rpm for 175, 140, and 100 knots IAS respectively.
With hydraulic pressure available from one windmill-
ing engine, and with extreme caution taken in rate of
control movement, the following can be completed
independently: extension of flaps partially or fully (if
left engine is windmilling); correction for slight tur¬
bulence; 30-degree bank turns; and flareout for land¬
ings. If both hydraulic systems fail, flight controls can
be operated with supplemental pump pressure if the
left system has not failed through loss of fluid. In ad¬
dition, all other hydraulic units can be operated, bur
discretion in their use should be exercised to avoid
lowering system pressure excessively. The supplemen¬
tal pump output is approximately equal to that of one
enginC'driven pump at idle rpm.
FLIGHT CONTROL SYSTEM
EMERGENCY OPERATION,
If the right or left hydraulic system fails, one 3000-psi
hydraulic system is available for basic flight control.
Normally, little difference will be noted with flight
under such conditions. This includes flight at maxi¬
mum level flight speed down to stall for the landing
configuration. Due to limited elevator deflection, avail¬
able load factor is lowered by approximately 0,3 "G.”
A limit in surface deflection occurs when there is a
balance in elevator power and airloads (limiting eleva¬
tor hinge movement). This means that the altitude lost
during recovery from a dive is greater with only one
hydraulic system operating. To be specific, maximum
load factor obtainable at 0.85 true Mach number and
10,000-foot altitude is approximately 2.0 "GY*; with
both systems operating about 2,3 "GY’ are available.
The limiting load factor, or "G” value, increases with
any one or a combination of the following; decrease
in Mach number, decrease in dynamic pressure, aft
movement of the airplane center of gravity, and a
decrease in horizontal stabilizer angle caused by man-
ufacturing tolerances. Under 0.80 Mach number, longi¬
tudinal control to limit load factor or airplane buffet
is available. Full basic control of the airplane is pos¬
sible in flight using the supplemental pump. Control
stick and rudder pedal actuating forces are comparable
3-27
Section til
T.O. 1F-89H-1
to those which occur with the normal system in opera¬
tion. The replenishing rate of the supplemental pump
is sufficient to maintain pressure during fast actuation
of the control surfaces, as would occur during flight in
turbulent air. Battery life when supporting the supple¬
mental pump and limited use of the radios is short. For
this reason it is suggested that the supplemental pump
be used only when absolutely necessary if 28-volt gen¬
erator power is unavailable. With only the hydraulic
pressure of one windmilling engine available, a safe
landing can be executed; however, it is necessary to
exercise extreme caution in the rate of control move¬
ment so as not to open the purge valve. The engine-
driven hydraulic pump replenishing rate at engine
windmill speeds is low; but full control deflections
applied at a slow rate, as necessary for a crosswind
landing, are possible. During flight in moderate to
heavy turbulence, basic stability should be depended
upon to a great extent for maintaining the selected
attitude.
With right hydraulic system pressure unavail¬
able, do not operate speed brakes, unless left
engine rpm is at least 85% or the supplemen¬
tal pump is operating. At lower rpm, the de¬
mand on the hydraulic system made by speed
brake operation will result in limited aileron
control unless supplemental pump pressure is
available.
SIDESLIP STABILITY AUGMENTER
EMERGENCY OPERATION.
If the sideslip stability augmenter system fails, causing
the airplane to oscillate violently, turn the sideslip
stability augmenter switch to PWR OFF. Without
stability augmentation, damping of the "Dutch Roll”
oscillation is extremely light under many flight con¬
ditions, but these oscillations can be controlled by the
pilot. Damping can be improved by descending to a
lower altitude.
If there is a malfunction of the electronic
control unit, making the stability augmenter
system inoperative with the rudder trim
switch in AUTO TRIM position, move the
switch to MANUAL TRIM. If the power
amplifier is still operative, the system will
continue to provide satisfactory damping of
"Dutch Roll” oscillations but may require
some manual adjustment of the rudder trim
knob.
WING FLAP SYSTEM
EMERGENCY OPERATION,
If the left hydraulic system fails through some cause
other than loss of hydraulic fluid, the wing flaps may
be lowered by normal procedures after actuating the
supplemental pump by depressing the nose wheel
steering button. If complete electrical failure occurs,
it is possible to lower the wing flaps if the left engine
is windmilling and the airspeed is below 150 knots
IAS; however, this procedure is not recommended be¬
cause of the long extension time and the possibility of
opening the purge valve with a resultant loss of the
complete system. In the event the pre-positioning
s P r * n 8 the flap handle has broken, the flaps may be
actuated by placing wing flap handle in desired posi¬
tion and moving the wing flap position indicator to
extend or retract flaps as needed. Considerable pressure
may be necessary to position the wing flap indicator
using this method.
SPEED BRAKE SYSTEM EMERGENCY
OPERATION,
The speed brakes cannot be operated if the left hy¬
draulic system fails; however, if the speed brakes are
open at the time of failure, they will float back to the
streamlined position when the speed brake lever is
placed at CLOSED. If speed brakes fail and remain in
the full open position, maximum power is required to
maintain level flight up to 15,000 feet.
LANDING GEAR SYSTEM EMERGENCY
OPERATION,
If the normal landing gear lowering procedure fails
to extend the gear to a safe condition, the pilot should
first try to determine what is causing the malfunction,
then execute the appropriate emergency procedure for
lowering the landing gear to a safe landing condition.
For example, the pilot can determine if there is flow
in the landing gear system by recycling the landing
gear lever from DOWN to UP and back to DOWN
while watching the left hydraulic system pressure gage
for fluctuations. If no fluctuations are indicated on the
pressure gage during the check, indicating no flow in
the landing gear system, it may be assumed that the
landing gear position 4 -way valve is stuck in the gear-
up position. If this occurs, the only way the gear can
be lowered is by reducing the left hydraulic system
pressure to zero. In order to accomplish this, the left
engine must be shut down, flaps partially lowered,
speed brakes partially opened, and then the flaps re¬
tracted at the same time the speed brakes are closed.
This will reduce the left hydraulic system pressure to
the point that the system purge valve (figure 1-26)
will open automatically (approximately 350 psi) and
reduce the system pressure to zero. However, the safety
3-28
T,0. 1F-89H-1
Section 111
relays circuit breaker in the radar observer's cockpit
must he pulled prior to flap and Speed brake operation
to disarm the left hydraulic supplemental pump. After
the system pressure has been reduced to zero, the land¬
ing gear emergency release handle may be pulled to
lower the gear.
CAUTION
• The left engine must remain inoperative after
the landing gear has been lowered by purg¬
ing the system pressure. If the engine is re¬
started, left hydraulic system pressure will be
restored to normal and the landing gear will
retract.
• When using the landing gear emergency re¬
lease handle, the pilot should make certain
the handle is pulled to its full limit of travel
(approximately 14 inches)* This will assure
that all landing gear up locks have been un¬
locked. The handle should then be returned
to its stowed position* Do not allow the
handle to whip back to its stowed position, as
damage to the cockpit equipment may result*
If any one or all of the landing gears fail to extend
after the landing gear lever is placed in the DOWN
position and the landing gear emergency release handle
has been pulled, the pilot should execute a coordinated
maneuver to pull positive "G*s,” This should be done
with the landing gear lever at DOWN and the emer¬
gency release handle pulled and held to its full limit
of travel. Care should be taken to avoid exceeding the
maximum allowable "GV* for the altitude at which the
maneuver is being executed*
GEAR FAILS TO EXTEND ON NORMAL
PROCEDURE.
L Airspeed—195 knots IAS or below, (P)
2, Landing gear lever—Check full DOWN, (P)
3, Left hydraulic system pressure gage—2000 psi. (P)
If pressure is below 2000 psi and time and condi¬
tions permit, allow pressure to build up.
Changed 13 February 1959
4* Landing gear emergency release handle—Pull to
full limit of travel (14 inches). (Allow at least 30 sec¬
onds for gear to extend,) (P)
i; CAUTION j;
The landing gear emergency release handle
should be guided back to its stow'ed position
to prevent the handle from whipping back
and causing damage to cockpit equipment.
5* Main landing gear—Check visually. (P)
6. Landing gear position indicators—Check for safe
indication* (P)
GEAR FAILS TO EXTEND ON EMERGENCY
PROCEDURE.
L Airspeed—195 knots IAS or below. (P)
2. Landing gear lever—Recycle, leave in DOWN
position. (P)
3- Left hydraulic system pressure gage—Check for
fluctuations. (P)
Note
If no fluctuations occur and the pilot is as¬
sured that no gears have moved, proceed with
the emergency procedure by purging the left
hydraulic system.
4, Left engine—Shut down, (P)
5. Safety relays circuit breaker—Puli. (RO)
6. Flap s—Lower partially, (P)
7, Speed brakes—Open partially, (P)
S, Plaps—Raise; speed brakes—Close simultaneously,
to open the left hydraulic system purge valve. (P)
9* Left hydraulic system pressure gage—Check for
0 psi* (P)
3-29
Section 111
T.O, 1F-89H-T
10, Emergency landing gear release handle—Pull {al¬
low at least 30 seconds for gear to extend). (P)
|; CAUTION I
The landing gear emergency release handle
should be guided back to its stowed position
to prevent the handle from whipping back
and causing damage to cockpit equipment.
Note
If gear fails to extend, continue with follow¬
ing procedures,
1L Emergency landing gear release handle-—Pull sec¬
ond time and hold at full limit of travel, (P)
12. Pull positive “G's.” (P)
13. Main landing gear—Check visually, (P)
14. Landing gear position indicators—Check for safe
landing gear indication, (P)
Note
After a prolonged flight at high altitude
(where temperature is low) emergency exten-
sion may be slower than normal,
GEAR FAILS TO EXTEND BECAUSE OF MECHANICAL
BINDING*
When the nose gear or main gear fails to extend be¬
cause of suspected mechanical binding, with hydraulic
pressure available, use the following procedures;
L Landing gear lever—DOWN. (P)
2, Landing gear emergency release "T JJ handle—
Pull to full limit of travel. (P)
3, Landing gear lever—UP, while maintaining ten¬
sion on 'T* handle in full out position, (P)
4, When gear has fully retracted, immediately place
landing gear lever DOWN, while maintaining tension
on T handle in full out position. After nose gear ex¬
tends, guide ”U* handle back to stowed position, (P)
5, Check gear down. (P)
Note
Lowering the landing gear by the emergency
procedure will not affect subsequent normal
operation. Each time the emergency gear ex¬
tension system is used, the pilot should report
it to ensure that the malfunction which neces¬
sitated the use of the emergency procedure is
corrected.
BRAKE SYSTEM EMERGENCY
OPERATION.
If the left hydraulic system fails, the brakes can still
be operated by the accumulator pressure. If necessary.
the radar observer can charge the accumulator by
placing the forward handpump selector valve (A) at
BRAKES, the rear valve <B) at NEUTRAL (see figure
4-8) and pumping the hydraulic handpump. A normal
ground roll stop can be made by using accumulator
pressure only, provided there is 3000 psi pressure in
the system. To stop the airplane using brake accumula¬
tor pressure, avoid too many applications which would
deplete hydraulic pressure. If wheel brakes fail to re¬
spond to brake pedal pressure, release brakes, immedi¬
ately turn the emergency airbrake handle to ON, then
operate the brakes as usual. When applying airbrakes
use caution as pedal resisting forces will be lighter than
normally experienced. If both emergency airbrake and
brake accumulator pressures are applied to the system
simultaneously, more pedal pressure than normal will
be required.
r
CAUTION
roN
Do not turn emergency airbrake handle to
ON while brakes are being applied; sudden
increase in braking efficiency may result in
a locked wheel and subsequent blowout.
Note
# The air bottle contains sufficient pressure for
three complete applications of the brakes.
• Brakes must be bled after using the emergency
airbrake system.
LOSS OF CANOPY,
If the canopy is lost, the airplane should immediately
be decelerated to 200 knots IAS or less. If no other
emergency exists, the emergency signal system should
be used, with prearranged signals, by the pilot and
radar observer for intercommunication.
Note
The following checklist is an abbreviated ver¬
sion of the procedures presented in the am¬
plified checklists of Section III. This abbre¬
viated checklist is arranged so that you may
remove it from your flight manual and insert
it into a flip pad for convenient use. It is
arranged so that each action is in sequence
with the amplified procedures given in Sec¬
tion III. Presentation of the abbreviated
checklist does not imply that you need not
read and thoroughly understand the amplified
version. To fly the airplane safely and effi¬
ciently you must know the reason why each
step is performed and why the steps occur in
certain sequence.
3-30
T.O. 1F-89H-1
Section IV
SECTION IV
Air for cabin air-conditioning and pressurizing and for
canopy defogging is taken from the Ilth stage of the
engine compressors. Ir then flows through a shutoff
valve in the supply duct to a bypass valve and refrig¬
eration unit, An electronic temperature-sensing system
automatically determines the settings of the bypass
valve* Cooled air from the refrigeration unit mixes in
the main duct with the hot air bypassing the unit and
flows through floor outlets into the cabin* (See figure
4-2.) A cabin temperature rheostat regulates the tem¬
perature of the air entering the cabin, and an automatic
pressure regulator controls the pressure. The cabin air-
conditioning system is controlled by 28-volt d-c power,
and the electronic temperature-sensing system is op¬
erated by 115-volt a-c power from the single-phase
inverter system.
Cabin Pressure Regulator.
The cabin is not pressurized below 12,500 feet. From
12,500 to 31,000 feet, the air pressure regulator main¬
tains the cabin pressure at the 12,500-foot altitude
pressure. Above 31,000 feet, the regulator normally
maintains a constant differential pressure of 5.00 psi.
For combat operation above 12,000 feet, an alternate
differential pressure of 2.75 psi can be selected so that
the drop in cabin pressure will not be explosive if the
cabin is suddenly depressurized. If the cabin pressure
regulator fails, a pressure-vacuum-relief valve relieves
HVAitmy emPMettr
T&BLI OF CONTENTS A
Cabin Air-Conditioning System ..* *
Canopy Defogging System.4-2
Anti-king Systems...- - *.4-5
Communication and Associated Electronic
Equipment ... . ....* 4-8
Lighting Equipment .. 4-20
Oxygen System ..* *.4-22
Autopilot . .. 4-26
Automatic Approach Equipment...4-29
Armament.....* * * -4-29
Optical Sighthead (Ml 69) .. 4-29
E-9 Fire Control System ... . 4-29
Single-Point Fueling System .. .4-29
Miscellaneous Equipment ................. .4-31
CABIN AIR-CONDITIONING SYSTEM*
excessive pressure. When the airplane dives to an alti¬
tude where the outside pressure is greater than that in
the cabin, the pressure-vacuum-relief valve opens to
equalize the pressure.
CABIN AIR SWITCH.
The 28-vok cabin air switch (figure 4-1) on the pilot's
instrument panel controls cabin air and pressure. When
the switch is at RAM & DUMP, ram air ventilates the
cabin, the engine compressor air is shut off, and the
cabin temperature control system is deenergized. When
the switch is at PRESS, the ram air is shut off, the
engine compressor air is turned on, and the cabin tem¬
perature control is energized.
CABIN DIFFERENTIAL PRESSURE SWITCH.
The cabin differential pressure selector switch (figure
4-1), a 28-volt d-c switch on the pilot's instrument
panel, provides a means of selecting either of two
available cabin pressures. For all normal operations this
switch should be at 5.00 psi so that from 12,500 feet
the cabin pressure regulator will maintain the cabin
pressure at the 12,500-foot level, and above 31,000 feet
Hf-4C
Section IV
T.O. 1F-89H-1
CABIN
AIN-CONDITIONING
CONTROL PANUS
Figure 4-1.
will maintain a constant differential pressure of 5.00
p$L For combat operations the switch should foe moved
to 2*75 psi to minimize any adverse effects if the cabin
is suddenly depressurized. (See figure 4-3.)
CABIN AIR TEMPERATURE SWITCH,
The cabin air temperature switch (figure 4-1) provides
a means for lowering or raising cockpit temperature
and is located on the pilot's aft miscellaneous panel.
The cabin air temperature switch operates on 28-volt
dc and has a center neutral position marked OFF;
other positions are AUTO, MOM. INCH, and MOM.
DECK. The switch is spring-loaded to OFF from the
latter two positions. When the switch is at AUTO,
the cabin temperature is maintained automatically
according to the setting of the cabin temperature
rheostat. When the switch is held at MOM. INCR
or MOM. DECR the cabin temperature rheostat is
cut out of the circuit and the cabin temperature
increases or decreases in proportion to the length
of time the switch is held. When the switch is released
to OFF, the cabin temperature is not automatically
controlled; the cooling unit bypass valve remains in
the position it is in and the temperature of the air
entering the cabin will remain constant if engine speed
and airplane altitude remain constant. The cabin air
temperature switch must be at AUTO when the pilot's
canopy defog knob is pulled all the way out; then a
sensing element, energized by the canopy defog knob,
can override the cabin temperature rheostat and main¬
tain a constant defogging air temperature of 79°C
(175°F).
CABIN TEMPERATURE RHEOSTAT.
The cabin temperature rheostat (figure 4-1) is a 28-volt
d-c knob on the pilot's aft miscellaneous panel. When
the cabin air temperature switch is at AUTO, the cabin
temperature rheostat automatically controls the tem¬
perature of the air in the cabin. The rheostat can be
rotated between COOLER and WARMER as desired
to control the temperature in the cabin* The rheostat
is out of the circuit when the cabin air temperature
switch is not at AUTO.
CABIN AIR-CONDITIONING SYSTEM
NORMAL OPERATION.
1* Cabin air switch-—PRESS.
2, Cabin air temperature switch—AUTO.
3- Cabin air temperature rheostat—As desired.
4* Cabin differential pressure switch—5.00 psi*
CABIN AIR-CONDITIONING SYSTEM
EMERGENCY OPERATION.
If the automatic temperature control fails, proceed as
follows:
1. Cabin air temperature switch—Hold momentarily
at MOM. INCR for warmer air or at MOM. DECR
for cooler air.
2. Wait a few minutes for change to become evident;
then repeat until desired temperature Is attained.
3. If this fails, place cabin air switch at RAM
DUMP.
CANOPY DEFOGGING SYSTEM.
Canopy defogging air is diverted from the cabin air-
conditioning floor outlets and released through ducts
along the canopy rail. The temperature of the air is
maintained at 79°C (175°F) by a separate temperature-
sensing unit. This sensing unit overrides the cabin
temperature rheostat if the cabin temperature switch
is set at AUTO and the pilot's canopy defog knob is
pulled all the way out.
CANOPY DEFOG KNOBS.
Two canopy defog knobs (figures X-ll and 4-7), one
on the pilot's center pedestal, and one on the left side
of the radar observer's cockpit, are provided for canopy
defogging. The defog knobs mechanically adjust valves
which divert the cabin air from the floor outlets to the
defogging ducts in the pilot's cockpit and torso com¬
fort outlets in the radar observer's cockpit. Each crew¬
member controls canopy defogging for his cockpit.
4-2
T.O. TF-89H-1
Section IV
Air-Conditioning System
CABIN AIR
PRESS.
£
AUTO
m
MOM MOM
INCR , DECR
From 28-roll d-c bus
From 115- roll a-c
single-phase
essen I ini bus
CABIN
A P-2./5
PSI
CABIN
A P-5.00
PSI
From cabin ) QOierboard
CABIN PRESSURE REGULATOR
DEFOG SWITCH ACTUATED BY
FINAL MOVEMENT OF PILOTS DEFOG KNOB
CONDITIONED AIR
RAM AIR
HOT AIR FROM COMPRESSOR
COLD AIR
ELECTRICAL ACTUATION
MECHANICAL ACTUATION
H606
Figure 4-2.
4-3
Section IV
T.O. TF-S9H-1
20
25
20
15
10
k
2,75
DIFFER
ENTIAL 1
PRESSU
RE ,
*
*
/
/
5.CX
3 DIFFEI
REIMTIAL
PRESSi
LIRE
/
1
1
#
*
CABIN
PRESSURE
SCHEDULES
H- 61 S
ngure 4-3.
but only the pilot’s defog knob can energize the sensing
element which overrides the cabin temperature rheo¬
stat and maintains the defogging air at 79°C (I75°F).
The pilot’s defog knob must be pulled all the way out
to energize the sensing element, and the cabin tempera¬
ture switch must be at AUTO to insure automatic
control of the air temperature.
CANOPY DEFOGGING SYSTEM OPERATION.
The canopy defogging system should always be used
immediately before and during descents.
Nate
The windshield and canopy defrost and defog
system should be operated at the highest
temperature possible (consistent with aircrew
comfort) during high altitude flights. This
high temperature will keep the transparent
surfaces preheated and will preclude forma¬
tion of frost and fog during descent.
If high humidities are known to exist at low altitudes
(dewpoint over 60° F) the defogging system should be
on for at least 30 minutes before descent to insure that
the canopy does not fog over at low altitude. The wind¬
shield is electrically heated to prevent defogging; wind¬
shield heat should be used at all times. The canopy
defogging system should be operated at the highest
possible temperature consistent with comfort during
high altitude flights, to preheat the canopy in order
to prevent the formation of frost and fog during
descents. When the canopy defogging system is used
at low altitude, correct procedure must be followed to
avoid overheating the canopy above its critical tem¬
perature of 8S.6°C to 933°C (190°F to 2Q0°F). At
these temperatures it softens and can fail under the
pressure loads that occur during certain flight con¬
ditions. The overheating itself does not permanently
damage the canopy, for when it cools back below the
critical temperature, it regains its original strength.
To obtain defogging air at the correct temperature,
the following steps should be performed in the order
given:
1. Cabin air pressure switch—PRESS,
2. Cabin air temperature switch—AUTO.
3. Pilot’s defog knob—PULL all the way out.
4. Radar observer’s defog knob—As desired.
Note
The radar observer should check with the
pilot to determine that all the pilot’s controls
affecting defogging are in their correct posi¬
tions before he pulls his defog control out, to
ensure controlled operation of the sysrem.
Step 3 fixes the automatic cabin air temperature con¬
trol at 79°C (I75°F) only if steps 1 and 2 have been
performed. Failure to perform step l will prevent any
control of temperature or pressure. Failure to perform
step 2 will leave the defogging air temperature un¬
controlled, affected only by compressor air temperature
and the position of the refrigeration unit bypass valve.
Failure to perform step 3 will leave defogging air
temperature uncontrolled, since only at the full out
position of the pilot’s defog knob will the defog tem¬
perature-sensing unit be energized to override the
cabin temperature rheostat when steps 1 and 2 have
been performed. If the cabin temperature switch is
held at MOM, INCR or MOM. DECK, the automatic
temperature control is overridden. If the pilot’s or radar
observer’s defogging knob is pulled out when the
4-4
Changed 13 February 1959
TO. 1F-S9H-1
Section IV
switch is held at MOM* INGR, air at full compressor
temperature is directed on the canopy and damage to
the canopy may result. If either knob is pulled out
when the switch is held at MOM* DEGR, air at the
lowest temperature available from the refrigeration
unit is directed on the canopy. The use of the defog
knobs at intermediate positions (not out far enough
to energize the automatic temperature control) when
the air-conditioning system is cooling the cockpits
will greatly increase cooling effectiveness, since air
from the defogging ducts will provide additional
cooling to the upper part of the body. Caution must
be exercised when the defog knobs are used in this
manner since damage to the canopy will result if heat¬
ing is turned on without returning the knobs to the
full in position.
ANTI-ICING SYSTEMS.
THERMAL AND ELECTRICAL ANTI-ICING SYSTEMS.
For the thermal anti-icing system, hot air is extracted
from the 11th stage of the engine compressor to anti-
ice the leading edge of the wings, empennage, and
engine air intake scoop. In normal operation, the hot
air maintains a predetermined leading edge skin tem¬
perature* The air passes through a pneumatic safety
valve and a modulating valve which is controlled
by the skin normal thermistors and the pressure con¬
trol If the normal thermistors fail to control the
modulating valve, and the surfaces of the leading
edges overheat, a skin overheat thermistor will close
the pneumatic safety valve to stop the flow of hot air
to the surfaces. When the temperature drops below a
predetermined value, the overheat thermistor will
reopen the safety valve until the surfaces again over¬
heat; then the cycle repeats. The engine inlet guide
vanes, bullet nose, island fairings, and forward frame
struts are heated by hot air bled directly from the 11th
stage duct whenever the anti-icing system is in opera¬
tion* Icing conditions are detected by means of a
pressure-sensing icing probe located in each engine
air inlet duct* When ice forms on either probe, a
28-volt d-c red warning light on the anti-icing control
panel illuminates, the engine screen normal controls
are overridden, and the engine screens are retracted.
When the airplane is parked with the power on and
the anti-icing switch is at OFF, the warning light will
come on and remain on, whether ice is present or not,
until the engines attain a speed of 62,5% rpm* Below
62*5% engine rpm the inlet air pressure is insufficient
to actuate the pressure switch* Operation of the ther¬
mal anti-icing system causes a rise in exhaust gas tem¬
perature, an increase in specific fuel consumption, and
a decrease in available thrust. The electrical controls
for the system operate on 28-volt dc. In the electrical
anti-icing systems, 28-volt d-c heating elements heat
the pitot tubes and engine icing probes. The fuel
tank vent heaters are energized by the 115/200-volt
alternator. The anti-icing switch controls the circuits
for all of these electrical heating units except the pitot
heaters. When the airplane is on the ground, a ground
safety switch on the main landing gear de-energizes all
circuits except the pitot hearing circuit.
WARNING
Do not use wing anti-icing during takeoff or
landing as maximum available thrust will be
reduced.
Note
The angle-of-attack probe heater is energized
at all times when the landing gear is re¬
tracted*
Anti-Icing Switch.
The 28-volt d-c anti-icing switch (figure 4-4) on the
anti-icing control panel controls the electrical circuits
of the thermal and electrical anti-icing systems* When
the red iighr warns that ice has formed on the icing
probes, the switch can he turned to TAKEOFF for
engine anti-icing or to FLIGHT for complete anti¬
icing* When the switch is at TAKEOFF or FLIGHT,
the electric heaters for the icing probes and fuel vents
operate when airborne; but while the airplane is on
the ground, the ground safety strut switch breaks
the circuits. When icing conditions no longer exist,
the anti-icing switch should be turned to OFF to de¬
energize all anti-icing circuits.
Wing Anti-Icing Override Switch
The 28-volt d-c wing anti-icing override switch (figure
4A) located on the anti-icing control panel, provides
manual control of the flow modulating valve if the
normal thermistor circuit fails* The switch has two
positions: NORMAL and EMER. When the switch is
placed at NORMAL, the modulating valve is controlled
automatically by the normal thermistors and the pres¬
sure control* When the switch is placed at EMER, the
modulating valve will open; however, if an over¬
pressure condition exists, the pressure control will
prevent the valve from opening regardless of switch
position. When the switch is at EMER, the overheat
thermistor will continue to control the pneumatic
safety valve.
Pitot Heat Switch.
Each pitot tube is heated by 28-volt d-c power. The
pitot heat switch (figure 4~4) on the anti-icing control
panel can be turned to OFF and ON to control the
4-5
Section IV
T.O. TF-89H-T
Pilot's h>p
ANTt-tem
comm PANELS
Figure 4-4 .
operation of the pitot heaters. The pitot heat switch
is not overridden by the ground safety switch and can
be turned to ON at any time.
Anti-Icing Warning Light.
When ice forms on the icing probes, the 28-volt d-c
anti-ice warning light (figure 4-4), located on the
anti-icing control panel, comes on to indicate that
the anti-icing system should be turned on. When the
anti-icing switch is placed at TAKEOFF or FLIGHT,
the light goes out and will not come on again while
the system is energized. When the anti-icing switch
is at TAKEOFF or FLIGHT the heating elements
for the icing probes are energized when ice forms
on the probes, and are automatically turned off when
the ice is melted.
Anti-Icing System Operation.
The following operating procedures are recommended
for use of the anti-icing system in conditions of known
icing or when indicated by the ice warning light.
Takeoff. Select TAKEOFF position of anti-icing
switch. This will retract the engine inlet screens and
provide hot air anti-icing of the engine forward frame
components.
9 Unless the anti-icing switch is placed at
TAKEOFF when taking off into icing con¬
ditions, the engine screens will remain ex¬
tended until the airplane leaves the ground.
In severe icing conditions the engine screens
may become iced within a few seconds, re¬
sulting in dangerous loss of power.
9 FLIGHT position of anti-icing switch is not
to be used on takeoff, because complete air¬
plane surface anti-icing increases the demand
on the compressor hot air bleed and causes a
much greater loss in thrust.
In Flight (Level Flight and Climb). Select FLIGHT
position of the anti-icing switch. This will retract
the engine inlet screens if screen switch is in EMER
EXTEN position, provide hot air anti-icing of the air¬
frame leading edge surfaces and engine forward frame
components, and provide electrical anti-icing of the
fuel vents.
Descent. In making a descent from altitude through
icing conditions, select FLIGHT position of anti-icing
switch, maintain a minimum of 83% engine rpm and
regulate airspeed and rate of descent as in normal
descent. If ice then accumulates (additional hot air is
required for anti-icing), increase the engine rpm with¬
out increasing airspeed.
Landing. Place the anti-icing switch in FLIGHT posi¬
tion before the final approach of a landing in icing
conditions with one or both engines operating to pro¬
vide ice protection for the wings and empennage. Use
of the anti-icing system affords protection against icing
conditions, but causes a decrease in available thrust.
If a go-around is necessary, the anti-icing switch may
remain in the FLIGHT position only if two engines
with maximum thrust and afterburning are available.
Place the anti-icing switch in TAKEOFF position dur¬
ing approach and landing under single-engine opera¬
tion in light or moderate icing conditions to provide
maximum thrust in case of a possible go-around.
Adequate ice protection is available from one engine;
however, available thrust may be dangerously reduced.
In most cases moderate icing of the airfoil leading
edges can be tolerated in preference to loss of engine
thrust. When a go-around is necessary with both
engines operating, but afterburners are inoperative, or
when a single-engine go-around is necessary, place the
4-6
T.O. 1 F-89H-1
Sedion IV
anti-icing switch ia the TAKEOFF position until a
safe go-around altitude is obtained. After reaching a
safe altitude, the anti-icing switch may be moved back
to FLIGHT position. In single-engine operation excess
thrust is low in landing and takeoff configurations.
Therefore, it is imperative that flaps and landing gear
are raised as soon as possible when making a single*
engine go-around.
Note
The hot air anti-icing systems use air from,
the engine compressor and thereby reduce
the available thrust, increase the specific fuel
consumption, and decrease the airspeed. The
anti-icing systems should therefore be turned
off when icing conditions no longer exist and
should not be turned on in the absence of
icing conditions,
LOW PRESSURE FUEL FILTER DE-ICING SYSTEM.
A low pressure fuel filter de-icing system is provided
for the engines. Alcohol can be injected into the fuel
filter to dissolve ice particles in the fuel filter and
engine fuel control. Fuel control icing will be evi¬
denced by a drop in rpm, by overspeeding, or lack of
throttle response in the affected engine. Overspeeding
or drop in rpm in excess of 2% while operating at
100% throttle setting can be construed as an icing con¬
dition. Alcohol from a 3.9 (US) gallon tank, located in
the right wing, affords approximately 3 minutes total
de-icing time. A 28-volt d-c pump supplies pressure
for operation of the low pressure fuel filter de-icing
system. Two solenoid valves, one for each engine, con¬
trol the flow of alcohol. Fuel filter or fuel control
icing is not necessarily associated with other icing
conditions, but will occur whenever water particles
exist in the fuel and temperature of the fuel falls
below 0°C (32°Fh
Low Pressure Fuel Filter De-Ice Switch.
A three-position 28-volt d-c switch, spring-loaded to
OFF (center) with other positions RIGHT and LEFT
(figure 4-4), is located on the anti-icing control panel.
This switch controls power to a 28-volt d-c motor-
driven de-icing pump, and opens either of two nor¬
mally closed solenoid valves in the lines from the pump
to the engine low pressure fuel filters. When engine
fuel control icing is indicated by variation in engine
rpm or lack of throttle response, the switch should be
held to the position representing the affected engine
(RIGHT or LEFT) until engine rpm ceases to fluctu¬
ate, indicating that fuel flow is back to normal. Nor¬
mal flow should resume in 30 seconds or less. When
the switch is released, the alcohol pump will stop
operating and the solenoid valve in the line to the
filter that was de-iced will return to its normally
closed position. The alcohol supply will allow ap¬
proximately 3 minutes of pump operation as the pump
delivery rate averages slightly more than 1 gallon
per minute.
Note
If foreign matter other than ice restricts the
flow of fuel through a filter, the correspond¬
ing engine will react as during icing. A filter
clogged by foreign matter will be indicated if
normal fuel flow does not resume after ap¬
proximately 30 seconds of de-icing operation.
This should not cause alarm. Before the fuel
pressure drop across the low pressure filter
becomes critical, a bypass valve will open
and fuel will be routed around the filter.
However, it is important to make sure that
the filter is cleaned immediately after com¬
pletion of flight.
RADOME ANTI-ICING SYSTEM.
The radome anti-icing system prevents ice, which
would cause radar interference, from forming on the
nose of the airplane. Anti-icing fluid is supplied from
a pressurized gallon tank to a nozzle which
atomizes and sprays the fluid over the exterior surface
of the radome. The tank is pressurized and the fluid is
atomized by air from the 11th stage engine mani¬
fold. The compressor air is controlled by a solenoid
valve actuated by switches in the pilot's and radar
observer's cockpits. To prevent thickening at low
temperatures, the fluid is maintained at about 40°F
by a thermostatically controlled heater in the tank.
For fluid specifications, see figure 1-45.
Radome Anti-Icing Switches.
The system is actuated by placing the pilot's anti-
icing switch (figure 4-4), located on the anti-icing
control panel, at FLIGHT; however, the radar ob¬
server is provided with a 28-volt d-c override switch
(figure 4-8), located on the right side of the cockpit,
which gives him complete control over the system.
The switch has three positions: NORMAL, OFF, and
EMER. If the override switch is at NORMAL, the
pilot's anti-icing switch controls the system. If the
override switch is at OFF, the system is off regardless
of the position of the pilot's anti-icing switch. If the
override switch is at EMER, the system is on regardless
of the position of the pilot’s anti-icing switch.
Note
0 The supply of anti-icing fluid will last ap¬
proximately 1 hour if used continuously. This
must be taken into consideration when plan¬
ning interceptions under icing conditions.
• When anti-icing fluid has been used, a nota¬
tion to this effect should be made in DD
Form 781.
WINDSHIELD HEAT SYSTEM,
The windshield is defrosted and de-iced by two trans¬
parent heat-conducting films within the windshield
glass. The defrost system utilizes 28-volt dc and 115-volt
4-7
Section IV
T.O, 1F-89H-I
single-phase inverter system ac for control and sensing
circuits, and alternating current from the 115-volt alter¬
nator or from the single-phase inverter system for
windshield heat. The temperature is automatically
controlled by heat-sensing elements and temperature
regulators*
WINDSHIELD DE-ICE AND DEFOG KNOB.
A 28-volt d-c rotary windshield de-ice and defog knob
(figure 4-4), on the anti-icing control panel has OFF,
NORMAL, and EMER positions to control the wind¬
shield defrost and de-ice circuits* For defrosting, the
knob is placed at NORMAL: full a-c power is supplied
to the inner heat-conducting film and medium a-c pow¬
er to the outer heat-conducting film* For de-icing, the
knob is placed at EMER, and full a-c power is supplied
to both heat-conducting films. The EMER position
should be used only for heavy icing conditions, and
the switch should be returned to NORMAL as soon as
possible* The EMER position should never be used
when the airplane is on the ground because the extreme
heat applied to the outer film could damage the wind¬
shield* Primary power for windshield heat is supplied
by the alternator; but if the alternator fails, the single¬
phase inverter system will supply power for the defrost¬
ing circuits.
CAUTION ;;
To prevent possible bubbling of the heat-
conducting film in the windshield glass,
leave the windshield de-ice and defog knob at
NORMAL for at least 1 minute before turning
it to EMER* Only in heavy icing conditions
should it be turned to EMER. Never operate
the system on EMER longer than necessary.
COMMUNICATION AND ASSOCIATED
ELECTRONIC EQUIPMENT.
Interphone System AN/AIC-10
The interphone system, operating on 28-volt dc, pro¬
vides the following facilities: speech communication
within the airplane with or without the use of micro¬
phone switches, communication beyond the airplane
by integration with its radio equipment, monitoring
of received signals either individually or simultane¬
ously, a call facility which permits transmission of
urgent communication to both headsets regardless of
individual control panel switch settings. On airplanes
modified in accordance with T.O. 1F-89-627, the
landing warning horn has been removed and replaced
with an audible warning signal unit. If the landing
gear has not extended and locked properly on air¬
planes so modified, a warning signal will be audible
over the pilot's headset. Operation and control of the
audible warning signal unit is the same as for the
landing gear warning horn which it replaces* Recep¬
tacles in the right wheel well and in the aft radio
and equipment section allow communication between
the airplane crew and the ground crew*
Interphone Control Panel AN/AIC-10*
An interphone control panel (figure 4*9) is located
on the right console in each cockpit. Each panel has a
volume control knob, five (toggle type) mixing switch¬
es, a* rotary selector switch, and an auxiliary listen
switch. The mixing switches, marked INTER, COMM,
MARKER, ADF, and VHF NAV, enable the operator
to monitor incoming signals from all five sources
(interphone, command, marker beacon, radio compass
or omnirange and localizer sets), or to select any com¬
bination* The rotary selector switch has positions
COMM, COMM-INTER, INTER, and CALL, starting
at the left and going clockwise. The switch's function
is conventional. For example: with the switch at
COMM-INTER or CALL, the microphone is open for
interphone communication, but with the switch at
either COMM or INTER, the operator must press a
microphone button to talk or transmit. The auxiliary
listen switch has NORMAL and AUX LISTEN posi¬
tions. The toggle is safetied at NORMAL (up)* When
the switch is moved to AUX LISTEN any incoming
signals bypass the interphone amplifier and come
into the headset at line level (unamplified).
ADF Filter Switch
The ADF filter switch panel (figure 1-13) is located
on the pilot's right console. The filter switch is con¬
ventional in function, and has VOICE, RANGE, and
BOTH positions to mix or filter voice and range
signals when the radio compass is receiving on loop
or antenna.
Pilot's Microphone Switches.
Two microphone switches (figures 1-7 and 1-28), one
located on the right engine throttle knob, and one on
the control stick grip, can be pressed to transfer the
microphone input from the interphone to the com¬
mand transmitter*
Radar Observer's Microphone Buttons.
One radar observer’s microphone switch (figure 4-7)
is located adjacent to the canopy defog knob and,
when pressed, transfers the microphone input from
the interphone to the command transmitter. A foot-
operated switch located on the floor under the radar
scope serves as a radio audio disconnect switch. When
pressed, it prevents all incoming radio signals from
reaching both the front and rear cockpits; however, the
radar observer can talk to the pilot on the interphone.
This arrangement permits the radar observer to shut
out temporarily any distracting radio noises while
concentrating on the radar scope.
4-8
T.O. 1F-89H-T
Section IV
Figure 4-5
radio compass
LOOP ANTENNA •
EMERGENCY SIGNAL LIGHT
Figure 4-7
Section IV
T.O. TF-89H-1
Figure 4-8,
4-12
TO, 1F-89H-1
Section IV
Interphone Operation,
L Filter switch—BOTH.
2. Interphone selector switch—COMM-INTER,
3. Interphone toggle switch—INTER.
4. Auxiliary listen switch—NORMAL,
5. Volume control knob—Adjust as desired*
R O s cockpit^ right side
' s right
console
INTERPHONE
CONTROL PANEL
Figure 4-9*
Note
The interphone set is in operation whenever
electrical power is on the airplane* unless
the interphone circuit breakers (on the radar
observer's circuit breaker panel) are pulled
out,
COMMAND RADIO AN/ARC-27,
The command radio set, operating on 28-volt dc, is
used for airplane-to-airplane and airplane-to-ground
communication. The range varies with the altitude
and atmospheric conditions. A UHF channel identifi¬
cation holder is located on the forward right sliding
canopy frame directly below the defog duct.
Command Radio Controls.
Control panels for the command radio (figure 4*10)
are on the pilot's right console and the radar ob¬
server's right console. Each control panel has a power
control switch, channel selector switch, volume con¬
trol knob, control-shift switch* and a green indicator
light. The control-shift switches transfer control of
the command radio to either cockpit, and the green
light comes on in the cockpit having control. To
transmit to the ground or to another airplane, a micro¬
phone switch must be depressed.
Command Radio Operation,
1. Power control switch—T/R. Allow equipment to
warm up for at least I minute,
2. Channel selector switch—Rotate to desired fre¬
quency channel. Set is now ready to transmit and
receive.
3. Power control switch—T/R T G RFC, if simul¬
taneous reception on guard-frequency channel and
another channel is desired.
4. Volume control knob—Adjust as desired,
5. Microphone button—Press to transmit.
6. Power control switch—OFF to turn set off.
9 When the command radio set has been turned
off* do not turn set on again for 1 minute.
Allowing the condensers to discharge pre¬
vents an excessive power surge.
® To avoid damage to the selector mechanism,
do not select another channel while set is in
midcycle.
Notfe
No transmission will be made on emergency
(distress) frequency channels except for emer¬
gency purposes. For test, demonstration, or
drill purposes, the radio equipment will be
operated in a shielded room to prevent trans¬
mission of messages that could be construed
as actual emergency messages.
R 0*s cockpit— right side
Pilot's right console
c
COMMAND RADIO
~ CONTROL PANU
Figure 4-10.
4-13
Section IV
T.O. 1F-89H-1
RO cockp it— right side
Pilots right console
RADIO COMPASS
CONTROL PANU
Figure 4-7F.
RADIO COMPASS AM/&RN-6<,
The radio compass, operating on 28-volt d-c power,
indicates the direction to any selected transmitting
station when the radio compass is set for homing op¬
eration of the loop antenna. The signal of this re¬
ceiver is fed to the No. 1 needle of each radio
magnetic indicator on the pilot’s and radar observer’s
instrument panels.
Radio Compass Controls,
Radio compass control panels (figure 4-11) are on
the right console of each cockpit. Each control panel
has a function switch, frequency band selector switch,
loop L-R switch, volume control knob, CW-voice
switch, and tuning crank. Either crewmember can
gain control of the radio compass by turning the
function switch to CONT.
Radio Compass Operation.
1. Function switch—CONT momentarily to gain
control; then turn to desired position. Allow at least
5 minutes for warmup.
2. Interphone selector switch—Any position.
3- Interphone ADF switch—ADR
4. Frequency band selector switch—Turn to desired
frequency.
5. Volume control knob—Adjust.
6. Function switch—OFF, to turn set off (both
cockpits).
Note
• Operation of the E-9 fire control system causes
mild to severe interference of the AN/ARN-6
radio compass, depending upon homing signal
strength and frequency selected!
• The function switch in either the pilot's or
radar observer's cockpit will turn the set off
only when the function switch in the other
cockpit is also in the OFF position.
WHF NAVIGATION SET AN/ARN-M.
This equipment receives visual omnirange, visual-
aural range, localizer, and communication signals in
the high-frequency range of 108.0 to 135.9 mega¬
cycles. It employs 280 channels spaced 100 kilocycles
apart, in the following categories:
FREQUENCY ALLOCATIONS
Frequency Band
in Megacycles Type of Service
108.0—11L9
108.3—110.3
111,0—111.9
Runway Localizer
Visual-Aural Range (VAR)
Weather Broadcasts
Pilot's right console
VHP NAVIGATION
CONTROL PANU
figure 4-12 .
4-14
T.O. 1F-89H-1
Section IV
RADIO MAGNETIC
INDICATOR
n-7ie
Figure 4- 13.
Frequency Band
in Megacycles Type of Service
112.0—117,9
118*0—121.9
122.0—135*9
Visual Omnirange (VOR)
Tower
General Communications
As the transmission in these bands is iine-of-sight,
reception varies from 3 miles unobstructed distance at
sea level, to approximately 100 miles at 10,000 feet,
and even greater distances at higher altitudes. The
dynamotor operates on 28-volt dc; the indicators
operate on 26-volt ac from the C-l amplifier. For
instructions covering use of this equipment for auto¬
pilot-controlled approach, see Automatic Approach
Equipment, this section,
VHF Navigation Set Controls.
The VHF navigation control panel (figure 4-12) on
the pilots right console has a power switch, a fre¬
quency selector knob, and a volume control knob.
The power switch is turned from OFF to ON
to put the set into operation. The outer ring of the
frequency selector dial rotates to show as a whole
number, megacycles from 108 to 135 in the top three
windows of the frequency selector dial. A center
knob selects intervals of hundred-kilocycles which
appear as decimal parts of a megacycle in the bottom
window of the dial,
VHF Navigation Set indicators.
Two indicators for this equipment are on the pilot's
instrument panel, A course indicator registers VOR,
VAR, localizer, and glide slope orientation, A radio
magnetic indicator combines the functions of a direc¬
tional indicator (slaved) with those of a dual radio
compass. A duplicate radio magnetic indicator is on
the radar observer's instrument panel.
Radio Magnetic Indicator. The radio magnetic in¬
dicator (figure 4-13) includes a rotating compass card
and two needles. The rotating card is coupled
to the gyrosyn compass system. The signals of the
radio compass are fed to the No. 1 needle; the signals
of the omnirange receiver are fed to the No. 2 needle
when the receiver is tuned to a VOR transmitter. The
angle between a needle and the index at the top of
the instrument face will give the relative bearing;
and the radio magnetic indicator will read, on the
card under the point of the needle, the actual mag¬
netic bearing to the station regardless of the heading
of the airplane. Since the card will hold to magnetic
north and the two needles will hold to the tuned
radio stations, the card and the needles will appear
to rotate as if fixed together whenever a tight turn
is made at some distance from the stations.
Course Indicator, The course indicator (figure
4-14) has a marker beacon indicator light in one corner
COURSE INDICATOR
H-72g
Figure 4-14.
4-15
Section IV
T.O* 1F-89H-1
and a course set knob in the opposite corner. On the
face of the instrument are: a course window which
displays the number of the omnirange radial set up
hy the knob; a sensing window which indicates
whether the radial course leads to or from the omni¬
range station; a relative heading needle which is
coupled to the gyrosyn compass system; a vertical slid¬
ing bar and a horizontal sliding bar. When the receiver
is tuned to a VOR station and the warning "off”
flags have retracted from the face of the instrument,
the instrument shows which of the 360 radials of the
omnirange station has been selected {course window),
whether that radial course leads to or from the station
(sensing window), whether the radial lies right or left
of the airplane (vertical bar indication), and whether
the airplane is headed right or left of the selected
course (relative heading needle). The horizontal bar
does not respond to VOR signals; but when a
glide-slope transmitter has been tuned in, the bar will
show the position of the airplane with respect to the
glide slope.
VHF Navigation Set Ground Check.
1. Single-phase inverter switch—NORMAL.
2. Three-phase inverter switch—MAIN.
3. Directional indicator slaving cutout switch—ON.
4. Interphone selector switch—Any position.
5. Interphone ADF switch—ADR
6. Interphone VHF switch—VHF NAV.
7. VHF power switch—ON.
8. VHF frequency selector knob—Set on frequency
of nearest omnirange station.
9. Radio compass function switch—CONT. When
reaction of meter indicates that control has been ob¬
tained, turn to COMP.
10. Course indicator—Check that warning "off" flag
has retracted from vertical bar after equipment has
had a 2 to 5 minute warmup,
1L Radio magnetic indicator—Note that compass
card reads the airplane heading and that No, 2 needle
swings to bearing of omnirange station.
12. Course set knob—Rotate to set bearing to VOR
station in course window. Note that vertical bar cen¬
ters, and that sensing window reads TO. Note that
relative heading needle is displaced to the same side
of the station as the airplane's heading. Rotate course
set knob to set up radials 7 degrees to right and 7
degrees to left, and note that vertical bar moves
promptly and smoothly to full deflection on appro¬
priate side. Continue rotating course set knob. When
difference exceeds 90 degrees, note that the vertical
bar crosses to the opposite side of instrument, and
sensing window shows FROM, When reciprocal radial
is reached, note that vertical bar comes to center.
13- VHF frequency selector knob—Tune to nearest
VAR or localizer transmitter, if one is within receiving
distance, and note that vertical bar makes correct
response.
14. Radio compass frequency band selector switch—
Tune to nearest suitable transmitter and note that
No. 1 needle of radio magnetic indicator swings to
proper bearing.
15. VHF power switch—OFF, to shut down receiver.
16* Radio compass function switch—OFF, to turn set
off*
VHF Navigation Set Operation.
1. VHF power switch—ON.
2. VHF frequency selector knob—Rotate inner and
outer ring of dial to select frequency,
3. VHF volume control knob—Adjust as desired*
4. VHF power switch—OFF, to turn set off.
VHF Navigation Set—Operation With VOR*
1. VHF power switch—ON.
2. VHF frequency selector knob—Set for desired
VOR station* Allow 2 minutes for warning "off”
flag to retract from vertical bar,
3* Course set knob—Rotate to center vertical bar.
Read radial in course window and identify it as course
to or from the station as indicated in sensing window.
Read relative heading needle to determine whether
aircraft is headed right or left of course. If reciprocal
is desired, rotate course set knob to add or subtract
180 degrees; read course and sensing as now indicated.
To fly on a radial other than the one the airplane
is on, set up desired radial in course window. Vertical
bar will then be deflected toward new radial. Fly
toward vertical bar to arrive at desired radial, then
turn onto course as bar centers. Adjust heading as
necessary to compensate for drift. As long as vertical
bar is centered, airplane is tracking along displayed
radial, regardless of heading. Relative heading needle
will indicate drift angle. When airplane crosses station
while tracking along displayed radial, sensing will
reverse with no changes in other indications of the
instrument. When airplane is not tracking along dis¬
played radial, vertical bar will be off center. In such
a case, bar will swing to opposite side when airplane
crosses displayed radial. To turn smoothly onto radial,
steer to hold point of relative heading needle on ver¬
tical bar until both are centered. Sensing will reverse
when airplane crosses the radial; that is, at 90 degrees
to displayed radial.
VHF Navigation Set—Operation With VAR.
1, VHF power switch—ON.
2. VHF frequency selector knob—Set to desired
VAR station. Allow 2 minutes for warning "off”
flag to retract from vertical bar.
4-16
T.O. 1F-89H-T
Section IV
3, Note deflection of vertical bar. If bar deflects to
left, airplane is in blue sector of range; if bar is to
right, airplane is in yellow sector. Consult airways
chart to identify sector.
Note
On VAR, the deflection of the vertical bar
does not in itself indicate the direction in
which to fly to get on course. It indicates
merely the color sector in which the airplane
is flying.
4, Identify signal in headphones as aural N or A,
and consult airways chart to determine whether station
is ahead or astern. If aural signals overlap to give a
continuous dash, airplane is on aural leg at right
angles to visual range,
5, Relative heading needle indicates heading relative
to course selected.
Note
Blue and yellow sectors are assigned to op¬
posite sides of the visual range in accordance
with the course defined by the airway. At
certain terminal airports, VAR is used in the
absence of a localizer. In such cases, the
sector orientation is the same as for an 1LS
localizer. That is, the blue sector is charted
on the right and the yellow sector is charted
on the left when the airplane is inbound on
final approach, regardless of the course de¬
fined by the beam,
VHF Navigation Set—Operation With Localizer.
1. VHF power switch—ON.
2. VHF frequency selector knob—Set to localizer
station. Allow 2 minutes for warning "off' flag to
retract from vertical bar.
3, Note deflection of vertical bar. If vertical bar
is deflected to left, airplane is in blue sector of
localizer range; if bar is to right, airplane is in yellow
sector. Blue sector of a localizer is always charted
to the right of the inbound course; therefore, a pilot
on final approach can center on the beam by flying
toward the bar.
4, Relative heading needle indicates required correc¬
tion angle.
VHF Navigation Set——Operation
For Communications.
The receiver can be tuned to the appropriate trans¬
mitter to receive weather broadcasts, tower instructions,
and general communications.
GLIDE-SLOPE RECEIVER AN/ARN-18.
The glide slope gives vertical guidance to a pilot mak¬
ing an instrument approach to an airport equipped
with a glide-slope transmitter. The receiver has no
separate control panel. It is operated and tuned by
the power switch and the frequency selector knob
(figure 4-12) on the VHF navigation control panel,
and its signals are fed automatically to the horizontal
bar of the course indicator. When the set is tuned
to a glide-slope transmitter and the signal is strong
enough to retract the warning "off 1 * flag from the
horizontal bar, the pilot merely keeps the horizontal
bar centered to follow the glide slope down to the
runway. In brief, centering the two crossbars of the
course indicator keeps the airplane on course and on
glide slope for an instrument approach under adverse
weather conditions. The set is powered by the single¬
phase inverter system.
MARKER BEACON RECEIVING SET AN/ARN-12.
The marker beacon receiving set gives visual and
aural coded signals whenever the airplane passes over
a marker beacon transmitter, thus enabling the pilot
to determine his exact position. The visual signal
is given by an amber light (figure 4-14) on the pilot's
course indicator, the aural signal through the inter¬
phone system whenever the interphone marker beacon
switch is at MARKER and the interphone selector
switch is on COMM-INTER. The set operates when¬
ever the 28-volt d-c bus is energized,
A-2 FLIGHT COMPUTER.
The A-2 flight computer electronically combines atti¬
tude, altitude, direction, and radio information on a
single instrument. The flight computer may be used
in flying a constant altitude compass course, in making
ground-controlled approaches, in making instrument
low approaches, and for go»around$. The radio rate unir
feeds into the computer a signal derived from the rate
of change of the localizer signal as the airplane nears
the runway, so that the pilot by keeping the vertical
bar centered, can fly the localizer beam heading with¬
out correcting for wind drift on the heading indicator.
This feature reduces the likelihood of over-correcting
for wind drift during the latter stages of a low ap¬
proach. The flight computer has a selector switch (fig¬
ure 4-15) and a flight computer indicator (figure 4-16)
on the pilot's instrument panel. The system is ener¬
gized whenever the airplane's electrical power is on
and the main or spare three-phase inverter is operating.
If the main and spare three-phase inverters fail, the
directional indicator on the flight computer will con¬
tinue to operate; however, the horizontal and vertical
bars will be inoperative.
Flight Computer Selector Switch.
The flight computer selector switch (figure 4-15) on
the pilot's instrument panel has LEFT, FLIGHT
INST, VOR-LOC RIGHT, and APPROACH positions.
When the selector switch is at FLIGHT INST, the
flight computer indicator is used as a flight instru¬
ment independent of radio signals. When the selector
4-17
Section IV
T.O. TF-89H-T
FLIGHT COMPUTER
SELECTOR SWITCH
H-73B
Figure 4-15 ,
switch is at any other position, radio signals are
relayed into the flight computer indicator for localizer,
approach, and landing purposes. When the selector
switch is on any position hut APPROACH and the
airplane is flying at the desired flight altitude, an
altitude control switch on the right side of the selector
switch can be turned to ON. Altitude control signals
will then be sent into the flight computer indicator
and any deviation in altitude will cause the horizontal
bar to move off zero. When the altitude control switch
is turned to ON, the pitch-trim knob on the flight
computer indicator becomes inoperative and the green
light in the lower left corner of the selector switch
goes out. When the selector switch is turned to
APPROACH, the green light comes on to indicate that
the altitude control has turned off automatically to
prevent conflicting signals from going into the flight
computer. When the selector switch is at APPROACH
and a go-around is necessary, the pilot can press the
altitude control switch and the horizontal bar will
move to indicate the optimum climbout angle.
Flight Computer Indicator.
The flight computer indicator (figure 4-16) centered
at the top of the pilot s instrument panel, has a course
dial, a directional indicator, and two crossbars, A course
set knob is on the lower left corner of the case and a
pitch-trim knob on the lower right corner. Turning
the course set knob rotates the course dial to bring
the desired track figure under the course index' at
the top of the instrument face. The directional indi¬
cator rotates simultaneously to repeat the reading of
the directional indicator of the gyrosyn compass system
so that the magnetic heading of the airplane can be
read continuously on the course dial under the heading
pointer. The vertical bar deflects to give an
appropriate "fly right"' or "fly left" indication. When
the pilot turns the airplane to zero the vertical ban the
directional Indicator follows the heading of the air¬
plane as it turns onto the new course. The vertical bar
will not go past zero unless the airplane is overcon¬
trolled in making the correction. When the airplane
is on the selected course, the directional indicator and
the vertical bars are centrally aligned with the course
index. Deviations in pitch, altitude, and glide slope
FLIGHT COMPUTER
INDICATOR
H-74B
Figure 4- 1 6 ,
signals cause the horizontal bar to move up or down.
The pitch-trim knob in the lower right corner adjusts
the horizontal bar to compensate for changes in air¬
plane pitch trim during flight. Clockwise rotation of
the pitch-trim knob causes the horizontal bar to give
a "fly up" indication.
Flight Computer Operation.
Starting and Ground Check.
1. Three-phase inverter switch—MAIN.
2. Directional indicator slaving cutout switch—ON,
3. Flight computer selector switch—FLIGHT INST.
4-18
T.O, 1F-89H-1
Section IV
4. Course set knob—Turn to make course dial read
the direction shown by the directional indicator* When
the flag disappears, indicating that the quick erector
has completed its cycle, the vertical crossbar should be
a pprox i m ately at zero and the direction a 1 indicator
should be aligned with the index*
5* Altitude control switch—ON* Horizontal bar
should nor move more than one needle width, if at
all. Green light should be off when altitude control
switch is ON*
6* Course set knob—Turn to rotate card to the right
and then ro the left; the vertical bar should signal
"fly left" and "fly right” respectively. Turn course
set knob to make course dial read the direction of the
directional indicator. Vertical bar should zero and
directional indicator should realign with the index.
7. Pitch-trim knob—Turn clockwise and counter¬
clockwise. Horizontal bar should move up and down
respectively.
8. VHP power switch—ON*
9. VHF frequency selector knob—Turn for proper
channel.
i; CAUTION
n-
Jlj
1 1
1 _
' I
Whenever sudden altitude changes in excess
of 500 feet are anticipated, the altitude con¬
trol switch should be turned OFF to prevent
damage to the altitude control unit.
10* Flight computer selector switch—APPROACH*
Vertical bar on the flight computer indicator should
move to left or right, depending on position of airplane
relative to the beam.
11* Altitude control switch—Push in to energize
go-around circuit. Horizontal bar should indicate "fly
up” and the orange flag should appear.
12. Flight computer selector switch — VOR-LOC
RIGHT, Orange flag and "fly up” indication should
disappear.
13- VHF power switch—OFF.
Flying Compass Course at Constant Altitude*
1* Selector switch—FLIGHT INST.
2* Course set knob—Rotate to bring desired track
figure on course dial under the course index. Vertical
bar will move ro right or left.
3. Vertical bar—Note deflection and fly to rezero
and to align directional indicator with course index*
4. Pitch-trim knob—Turn to zero horizontal bar
at desired airplane pitch attitude,
5. Altitude control switch — ON when airplane
reaches desired altitude. Green light on the selector
switch should go out.
6. Fly airplane to keep horizontal and vertical bars
zeroed at all times.
Note
When changes in airplane trim are required,
turn the altitude control switch to OFF until
the new attitude is established.
IFF AN/APX-6A*
The purpose of the IFF equipment is to identify as
friendly the airplane in which it is installed when
challenged by an interrogator-responsor associated
with friendly radars* When a radar target is accom¬
panied by a proper IFF reply, that target is considered
friendly. This system operates on 28-volt dc from the
primary bus and 115-volt ac from the auxiliary a-c
bus,
IFF Controls.
The master control knob and mode selector switches
are on the IFF control panel (figure 4-17) located on
the pilot's right console.
IFF Normal Operation.
Turn the IFF equipment on by placing the master
control knob at STBY, The tactical situation or the
communications officer will determine the ultimate
4-19
Section IV
T O. T F-89H-I
Pilot's right console
Iff CONTROL PANEL
H-75S
Figure 4-17,
position of the master control knob and mode switches
for each mission. To turn the equipment off, place the
master control knob at OFF*
IFF Emergency Operation*
For emergency operation, press dial stop and turn the
master control knob to EMERGENCY, On airplanes
modified in accordance with T.O. 1F-89-604 an ejection
notification switch has been installed on each crew¬
members ejection scat. When either pilot's or radar
observer's seat is ejected from the aircraft, the ejection
notification switch automatically actuates the emer¬
gency mode of the AN/APX-6 IFF system,
LIGHTING EQUIPMENT*
EXTERIOR LIGHTING*
Positron Lights and Control Switches*
The position lights are conventional in color and
arrangement and operate on 28-volt dc The position
light switch (figure 4-18) on the pilot's aft miscel¬
laneous control panel has STEADY, OFF, and FLASH
positions for controlling the operation of the lights.
A switch to the right of the position light switch has
DIM or BRIGHT positions to determine the intensity
of the position lights. A flasher unit flashes the posi¬
tion lights at 40 cycles per minute; if the flasher
unit fails, steady operation of the lights is automatic.
The circuit breaker for the position lights is on the
radar observer's circuit breaker panel,
Landing-Taxi Light and Control Switches.
The single retractable light, located on the under side
of the fuselage nose section just forward of the nose-
wheel, serves for both landing and taxiing. The light
is controlled by two switches (figure 4-18) on the left
vertical console; an extension-retraction switch with
EX I END, RETRACT, and OFF positions; and a light
switch with LANDING, TAXI, and OFF positions.
The light is extended or retracted by placing the ex¬
tension-retraction switch at EXTEND or RETRACT
The light may be stopped in any position along the
arc of travel by placing the switch at OFF, Extension
or retraction takes about 10 seconds. Limit switches
automatically stop the extension-retraction motor when
the light is fully extended or retracted. The light is
turned on and off by the light switch. When the
Pitot's miscellaneous control panel
POSITION LI6NTS
CONTROL PANU
LIGHTING
CONTROL PANELS
/
Pilot‘s left ^ /
vortical console
LANDING-TAX!
L/ONT SWITCHES
H-7'68
Figure 4-18,
4-20
TO, 1M9H-1
Section IV
switch is placed at LANDING (with the extension-
retraction switch at EXTEND)* the light burns at
maximum intensity and is positioned at the correct
angle for landing or takeoff. When the switch is
placed at TAXI with the extension-retraction switch
at EXTEND* the light is positioned at the correct
angle for taxiing (about 7 degrees higher beam than
for landing) and the light beam widens and dims. The
light can be turned on before extension if necessary
so that the heat generated by the filament will de-ice
the light assembly. After retraction, the light must
be turned off by the light switch. The light and con¬
trol switches are powered by the 28-volt d-c primary
bus.
!; CAUTION ;;
I }«########+######### I
The landing-taxi light generates intense heat
which may damage the light; therefore, do not
use the light in the landing position longer
than necessary. On the ground do not use the
light in either position when the airplane is
not moving.
Note
When changing from one position of the
light (taxi or landing) to the other, the ex¬
tension-retraction switch must be placed at
EXTEND; otherwise the extension-retrac¬
tion motor will not operate and the light
will remain in the original position,
INTERIOR LIGHTING.
Pilot's Cockpit Lighting.
Red floodlights, operating on 28-volt dc, light the
pilot's instrument panel and cockpit area. Two are on
the movable section of the instrument panel glare
shield; others are on the left and right sides of
the cockpit structure. Three red floodlights, spaced
evenly below the rail on each side of the cockpit,
light the pilot's consoles. Red bulbs under individual
ring-type, hinged lighting shields illuminate the flight
instruments. The engine instruments and the fuel quan¬
tity gages are lighted by red floodlights. Indirect
plastic panel lighting is used for all other panels,
control position indicators, and markings. A C-4
cockpit light with a removable red filter can be
swiveled or removed from the mount. Two rheostat-
controlled thunderstorm lights, operating on 28-volt
dc, are provided to counteract temporary blindness
when eyes, adapted to the dark, are exposed to light¬
ning flashes. These lights also provide interior illumi¬
nation required for high altitude daytime flying- They
consist of two white floodlights mounted one on each
side of the pilot's cockpit approximately 4 inches above
the left and right consoles and aligned so that their
light beams converge on the lower center of the in¬
strument panel. On the cockpit lighting control panel
individual rheostats are provided to control the opera¬
tion and intensity of the floodlights, instrument ring
lights, and indirect lighting. A warning light dimming
switch, located on the same panel, can be used to dim
the warning lights during night operations. All light¬
ing circuits for the pilot’s cockpit are protected by
circuit breakers on the pilot's circuit breaker panel. A
stowage case for spare bulbs (figure 1-13) is attached
to the bulkhead aft of the right console.
Pilot's Cockpit Lighting Rheostats. Seven 28-volt d-c
rheostats (figure 1-13), located outboard of the pilot's
right console, rotate from OEF to DIM to BRIGHT
to control the pilot's cockpit lighting circuits. The
first rheostat at the forward end of the pilot's cockpit
lighting control panel controls the plastic panel lights;
the second, the console lights; the third controls the
instrument panel floodlighting, and the fourth, the
console floodlighting. The fifth rheostat controls the
lighting for the engine instruments; the sixth rheostat
controls the lighting for the flight instruments. The
thunderstorm light rheostat, mounted outboard of the
pilot's right verticle console, rotates from OFF to DIM
to BRIGHT to control both thunderstorm lights.
Warning Lights Dimming Switch. A 28-volt d-c warn¬
ing light dimming switch (figure 1-13), located on the
pilot’s cockpit lighting control panel, provides a means
of dimming, during night flying, all warning and
indicator lights except the fire and over-heat warning
lights, oxygen indicator, and inverter failure warning
lights. The switch has DIM and BRIGHT positions
and is spring-loaded to an unmarked NEUTRAL
position. When the switch is momentarily placed at
either position, the warning light intensity will be that
of the selected position. The switch is interconnected
with the flight instrument lighting rheostat and will
not control warning light brightness if the rheostat is
at OFF.
Radar Observer’s Cockpit Lighting.
Two 28-volt d-c red floodlights, mounted under the
radar observer's glare shield, light the cockpit area.
Two red bulbs under individual ring-type lighting
shields illuminate each instrument on the instrument
panel. The shields are hinged to permit replacement
of the bulbs. All other panels have indirect or flood
lamp lighting. A C-4 cockpit light can be swiveled
or removed from its mount for either red or white
lighting. Four rheostats control the operation and
intensity of the instrument and circuit breaker flood¬
lights, instrument indirect lights, console plastic panels,
and console floodlights. The circuit breakers for the
lights are on the radar observer's circuit breaker panel.
Radar Observer’s Cockpit Lighting Rheostats.
Four 28-volt d-c rheostats (figure 4-7) on the interior
lights control panel located on left side of the cockpit.
4-21
Section IV
T.O. 1F-89H-1
rotate from OFF to DIM to BRIGHT. The rheostat at
the top left controls the plastic panels; the top right
rheostat controls the instrument and circuit breaker
floodlights; the bottom left rheostat controls the con¬
sole floodlights; and the bottom right rheostat controls
the instrument indirect lights.
Pilot's and Radar Observer's Cockpit Lights.
A removable 28-volt d-c swivel mounted G*4 cockpit
light with a red filter is mounted in each cockpit. The
pilot’s light is stowed on the left console with an
alternate socket on the left windshield frame. The
radar observer's light is stowed above the right con¬
sole. A knob near the back of the light case turns
the lamp on and off and controls its intensity. A
white spring-loaded button on the back of the case
can be pressed for momentary lighting. A small knob
extending through a groove on the side of the case can
be moved for spot- or floodlighting; tightening the
knob screw locks the shield in any position. The red
filter can be removed, if white light is desired.
OXYGEN SYSTEM,
The airplane is equipped with a gaseous oxygen
system having operating pressure of 4i)0 to 450 psi.
The oxygen is carried in four oxygen cylinders which
are check-valved and installed in the aft fuselage for
combat safety. Two of the cylinders supply oxygen to
the pilot, and two supply the radar observer. Fach
crewmember's supply system is kept separate by the
seated check valves at the filling manifold. When the
check valves are unseated during filling, interflow
between the four oxygen cylinders supplying the pilot
and radar observer occurs. However, loss of pressure
in one cylinder will result in the check valves being
si
: },* ;
GAGE PRESSURE
1 PSI }
400
350
300
250
200
150
lOO
35.000
Sf ABOVE
5.7
4.9
4.0
3,2
2.4
1.6
0,8
5.7
4,9
4 0
3.2
2,4
1.6
0,8
30.000
4.1
3.5
2.9
2.3
1.8
1.2
0.6
4.2
3.6
3.0
2.4
T.8
1.2
0,6
25,000
3.2
2,7
2.3
1.0
1,4
0.9
0.5
4.0
3.4
2.8
2,3
1.7
1 1
0.6
20,000
2 4
2.T
1.7
1.4
1.0
0,7
0,3
4.5
3.8
3.2
2.6
1.9
1.3
0,6
15,000
2.0
1.7
1.4
1.1
0.9
0.6
0.3
5 4
4,7
3.9
3.1
2.3
1,6
0,8
10,000
2.0
1.7
1.4
1.1
0.9
0,6
0 3
5.4
4 7
3.9
3,1
2.3
1.6
0.8
below
100
faj < Z
C eg
a*
L53 e o'
sa <■*
fas 8 r
co r*
g e
a z
ROLl) FACE (UPPER) FIGURES INDICATE 1)1 LI TER LEVER "100%.”
LIGHT FACE (LOWER) FIGURES INDICATE DILUTEE LEVER "NORMAL.
OmtH DURATION NOONS CNART
CYLINDERS: 4 TYPE F2
CREW: 2
The above figures apply whether one or tiro crew- f
members are using oxygen, as each members system
is separate from the others. .
. H-78C /
4-22
TO. 1F-89H-1
Section IV
seated in the three remaining cylinders. The other tank
in the system of the rank losing pressure is the only
remaining source of oxygen to the crewmember being
supplied by the system containing the damaged tank.
On each crewmember's right console is an oxygen
regulator panel which contains the oxygen system con¬
trols. A pressure-demand oxygen mask should be used
with this system. The approximate duration of the
oxygen supply at various altitudes is given in figure
4-19.
OXYGEN REGULATOR.
A diluter-demand oxygen regulator control panel (fig¬
ure 4-20} with a pressure gage and flow indicator is lo¬
cated on the right console of each cockpit. From sea
level to 30,000 feet (cabin altitude) the regulator auto¬
matically varies the ratio of oxygen to air to supply the
proper mixture to the crew. Above 30,000 feet (cabin
altitude) the regulator delivers pure oxygen at maxi¬
mum pressure. A relief valve in the regulator prevents
excessive pressure in the oxygen mask.
Regulator Supply Lever.
The oxygen supply lever (figure 4-20) on the regula¬
tor panel controls oxygen flow to the regulator. On
airplanes equipped with the D-2 oxygen regulator,
the shutoff valve is safetywired in the ON position
in the pilot's cockpit to prevent accidental closing off
of the oxygen supply during use at altitude. The
radar observer’s oxygen supply lever should be turned
OFF whenever the radar observer's regulator is not
being used. If it is left at ON, oxygen will be lost.
Note
Because of the automatic pressure-breathing
feature of the regulator, a continuous flow
of oxygen at altitude will result if the regu¬
lator is not being used and the supply lever
is left at ON. This condition causes a rapid
loss of oxygen at altitude.
Regulator Diluter Lever*
The diluter lever (figure 4-20) on the oxygen regu¬
lator panel has two positions: NORMAL OXYGEN
and 100% OXYGEN. When the lever is at NORMAL
OXYGEN, the regulator automatically varies the ratio
of oxygen to air and supplies the proper mixture to
the crew from sea level to 30,000 feet. Above 30,000
feet the regulator delivers pure oxygen. At any alti¬
tude, the diluter lever can be turned to 100% OXYGEN
if pure oxygen is desired for emergencies.
Regulator Emergency Lever,
The emergency toggle lever (figure 4-20) should re¬
main in the center position at all times, unless an
unscheduled pressure increase is required. Moving the
toggle lever either way from its center position pro¬
vides continuous positive pressure to the mask for
on airplanes modified in
ACCORDANCE WITH T.O. 15X6-5-2-511
ON AIRPLANES MODIFIED IN
ACCORDANCE WITH T.O. IF-1-533
li O s rorknif- right side
Sk k* m s r *$* 1 rouSi ^ e
(-StSw r af l P kwi **.*)
REGULATOR PANEL
H 79B
Figure 4-20.
4-23
Section IV
T.O. 1F-89H-1
emergency use. When the lever is depressed in the
center position, it provides positive pressure to test the
mask for leaks. Normally the lever should remain at
the center OFF position.
[
CAUTION
When positive pressures are required, it is
mandatory that the oxygen mask be well fitted
to the face. Unless special precautions are
taken to ensure no leakage, continued use of
positive pressure under these conditions will
result in rapid depletion of the oxygen supply.
Regulator Warning System Switch
and Indicator Lights (Some Airplanes).
The warning system switch (figure 4-20) on the oxygen
regulator panel can be placed at ON or OFF to con¬
trol the oxygen warning lights. Two warning lights
(figures 1-8 and 4-6 ) are on the instrument panels in
the pilot's and radar observer's cockpits. One light
indicates breathing in the pilot's mask; the other
indicates breathing in the radar observer's mask.
When the warning system switch is ON, the light
dims when oxygen is being used and glows brightly
when oxygen is not being used. On airplanes modified
in accordance with T.O. IF-1-533 the warning system
switch is placed in the OFF position, the lamps are
removed from the warning lights, and the warning
system is deactivated.
Note
At flight altitudes below 10,000 feet with
the oxygen regulator operating and the mask
being used, the lights may blink brightly.
Since this can happen only at low altitudes,
it should not cause undue concern. The oxy¬
gen regulator warning circuit may be turned
off below 10,000-foot flight altitude since the
possibility of hypoxia is critical only at higher
altitudes.
Oxygen System Pressure Gage
And Flow Indicator.
A combination pressure gage and flow indicator (fig¬
ure 4-20) on the oxygen regulator panel registers the
pressure of the oxygen supply on the upper half of the
dial. In the lower half of the dial, the slots in the
flow indicator are luminous when oxygen is flowing
through the regulator, dull black when It is not.
Note
As an airplane ascends to higher altitudes
where the temperature normally is quite low,
the oxygen cylinders become chilled. As the
cylinders grow colder, the oxygen gage pres¬
sure is reduced, sometimes quite rapidly. With
a 100°F decrease in cylinder temperature, the
gage pressure can be expected to drop 20 per¬
cent. This rapid fall in pressure is occasionally
a cause for unnecessary alarm. All the oxygen
is still there, and as the airplane descends to
warmer altitudes, the pressure will tend to rise
again, so that the rate of oxygen usage may
appear to be slower than normal. A rapid fall
in oxygen pressure while the airplane is in
level flight or while it is descending, is
not ordinarily due to falling temperature, of
course. When this happens, leakage or loss of
oxygen must be suspected.
OXYGEN SYSTEM PREFLIGHT CHECK,
1. Mask male connector attachment strap—Fasten
to the parachute chest strap by routing the connector
strap up under the chest strap as close to the center as
possible, then down in front of the chest strap, and
around again, then snap it to the connector.
2. Mask-to-regulator tubing female disconnect—Con¬
nect to the mask male connector, listen for the click and
see that the sealing gasket is only half exposed.
3. Alligator clip—Attach to the end of the mask
male connector strap. (See figure 4-21.)
4. Oxygen regulator—Check with diluter valve first
at NORMAL OXYGEN and then at 100% OXYGEN
by blowing gently into the end of the regulator
tubing as during normal exhalation. If there is a
resistance to blowing, the system is satisfactory. Little
or no resistance to blowing indicates a leak or mal¬
function,
5. Oxygen warning light switches—ON (some air¬
planes). Warning light should emit a bright (steady or
blinking) light. Move emergency toggle from center
(OFF) to LEFT or RIGHT position. The warning light
should change from a bright light to a filament glow
and back to a bright light. Return emergency toggle
to center (OFF) position.
Note
9 Items pertaining to the oxygen warning sys¬
tem apply only to airplanes not modified in
accordance with T.O. IF-1-53 3-
• Conduct the following check with regulator
supply valve at ON, oxygen mask connected
to regulator, diluter lever at 100% OXYGEN,
and normal breathing.
6. Blinker—Observe for proper operation. Warning
light (some airplanes) should change from bright to a
dim filament glow.
7. Emergency toggle—Deflea to RIGHT or LEFT.
A positive pressure should be supplied to mask. Hold
breath to determine if there is leakage around mask.
Return emergency toggle to center (OFF) position;
positive pressure should cease.
8. Diluter lever—Return to NORMAL OXYGEN.
4*24
T.O, IF-89H-1
Section IV
WARNING
Failure to double-loop tiedown strap around
chest strap may permit tiedown strap to slip
into and open the chest strap snap during
ejection.
Fasten the attachment strap on the mask male
connector to the parachute chest strap by
routing the connector strap up under the
chest strap as rinse to the center as possible , then
down in front of the chest strap , and around
again ; then snap it to the connector.
PARACHUTE CHEST STRAP
Connect the m as k-t ^regulator tubing female
disconnect to the mask male connector ,
frstpri for the click and see that the
sealing gasket is tmly half exposeil.
MASK
MALE
CONNECTOR
ALLIGATOR CLJP
Attach the alligator clip to the FEMALE DISCONNECT
end of the mash male connector strap.
H-aoc
Figure 4-2 1.
OXYGEN SYSTEM NORMAL OPERATION.
1. Regulator diiuter iever—NORMAL OXYGEN.
2. Regulator supply lever—ON.
3* Regulator warning system switch—ON (on air¬
planes nor modified in accordance with T.O. 1F-1-533).
OXYGEN SYSTEM EMERGENCY OPERATION.
If either of the crew detects symptoms of hypoxia, or
if smoke or fuel fumes enter the cabin:
1. Regulator diiuter lever“I00% OXYGEN,
2, Regulator emergency lever—Push IN and hold
momentarily to clear mask.
emergency lever IN. The other member's sup¬
ply will nor be affected since the systems are
independent.
3. Oxygen diiuter lever~--NORMAL OXYGEN after
the emergency. If the oxygen regulator fails or if the
mask develops a severe leak, push the regulator emer¬
gency lever to RIGHT or LEFT. If necessary, pull
the cord of the bailout bottle.
Note
The duration of the oxygen supply for the
pilot or radar observer is reduced when either
turns to 100% OXYGEN or holds the oxygen
If either crewmember uses his bailout bottle,
the airplane must immediately be flown to
an altitude at which the crew does not re¬
quire oxygen.
4-25
Section IV
T,0. 1F-89H-1
AUTOPILOT.
An E-ll ail-electric autopilot, powered by the 28-volt
d-e main bus and the 115-volt a-c single-phase essential
inverter bus, can be used to fly the airplane in straight
and level flight, coordinated turns, climbs and descents
with or without maneuvering turns, and instrument
approaches. It can be engaged, without producing
abrupt changes in control or airplane attitude, at any
time the airplane is being flown within autopilot
engaging limits. This is due to ao automatic synchro¬
nization system which keeps the autopilot bridge cir¬
cuits electrically in trim during the time the autopilot
is disengaged. The autopilot can be manually over¬
powered with the control stick and rudder pedals.
Autopilot controls (figure 4-22) are grouped in mo
panels; the function selector and the flight controller,
both located on the pilot’s right console. Autopilot
controlled flight at constant altitude is made possible
AUTOMATIC PtlOT
CONTROL PANU
H4TB
Figure 4-22.
by an altitude control feature which derives its signal
from a sensitive aneroid. Signals from the gyrosyn
compass system provide a directional reference when
the manual turn control is not being used. A vertical
gyro provides a reference for measuring airplane dis¬
placement in the roll and pitch axes. Three rate gyros
(yaw, roll, and pitch) supply signals proportional to
rate of change of airplane displacement. When these
signals are added algebraically to the signal provided
by the vertical gyro, the result is a smooth coordination
of the flight controls in both the starting of maneuvers
and the return to straight and level flight. An auto¬
matic trim feature trims the elevator force-producing
mechanism while the autopilot is engaged, so that at
any time the autopilot is disengaged, control stick
forces will be at a minimum. A localizer and glide-slope
coupler provide means for automatic flight control
during the approach and glide-slope phases of instru¬
ment landing procedure. After the autopilot is en¬
gaged it will control the airplane through a maximum
of 60 degrees of roll and 50 degrees of pitch in either
direction from the horizontal. The engaging limits
are 50 degrees of pitch, 29 degrees of roll, and 10
degrees of yaw. The elevator servo contains a slip
clutch which limits servo output to 13 pounds of
stick force in the pitch axis. This limits "G's” during
autopilot controlled flight to a safe value for all flight
conditions. Autopilot aileron deflection is limited
to 5 degrees. When the autopilot is engaged, the air¬
plane displacement signals to the sideslip stability acg-
menier are interrupted, but the latter system remains in
standby status and will resume its stabilizing function
the instant that the autopilot is disengaged,
POWER SWITCH.
An autopilot power switch (figure 4-22), located on
the function selector panel, controls the electrical
power supply to the autopilot system. When the switch
is placed at ON, power is supplied to the autopilot
system. When the switch is placed at OFF, all elec¬
trical power to the autopilot is disconnected and the
engaging switch, if at ENGAGE, snaps to OFF.
ENGAGING SWITCH.
The autopilot engaging switch (figure 4-22) is located
on the flight controller panel and has an ENGAGE and
an unmarked OFF position. The switch is solenoid-held
and will remain in the ENGAGE position only when
the following conditions have been met: the autopilot
circuit breakers are IN, the power switch has been at
ON for 90 seconds or more, the turn knob on the flight
controller is in detent, and the airplane is in an atti¬
tude within autopilot engaging limits. When the en¬
gaging switch is placed at OFF the autopilot disen¬
gages. The switch will snap to the OFF position if
the power switch is turned to OFF or if the pilot's
emergency disconnect switch is used to disengage the
autopilot.
4-26
T.O. 1F-89H-1
Section tV
EMERGENCY DISCONNECT SWITCH.
A 28-volt d-c, spring-loaded lever-type emergency
disconnect switch (figure 1-28), located on the control
stick grip, provides a means of instantaneous auto¬
pilot disengagement. If the autopilot is engaged,
squeezing the emergency disconnect switch will dis¬
engage the autopilot and cause the engaging switch
to snap to OFF. The autopilot power switch will re¬
main at ON until manually moved to OFF. When
the emergency disconnect switch is used to disen¬
gage the autopilot, any of the solenoid-held coupler
switches that may be at ON at the time (altitude
switch, localizer switch, or approach switch) will
snap to OFF. Squeezing the emergency disconnect
switch also will reset the autopilot engaging circuit.
TURN KNOB,
A turn knob (figure 4-22), located on the flight con¬
troller panel, provides a means of making coordinated
turns with the autopilot. The knob normally rests in
a neutral detent (knob pointing forward). When the
knob is in this position, directional signals from the
airplane's gyrosyn compass system provide the auto¬
pilot with a heading or directional reference. Moving
the turn control knob to the right or left out of the
detent will result in an autopilot controlled coordi¬
nated turn in the direction that the knob is turned
and at a bank angle proportional to the amount the
knob is turned, up to a maximum bank angle of 60
degrees. When the turn knob is returned to the neutral
detent, the airplane will roll smoothly out of the turn
and continue to fly at the new compass heading. The
autopilot will not engage with the turn knob out of
detent.
HEADING TRIM INDICATOR AND KNOB,
The heading trim indicator and knob (figure 4-22)
on the flight controller are used to indicate and
correct heading mistrim during autopilot controlled
flight. To correct heading mistrim, rotate the heading
knob in the direction of needle deflection: clockwise
for right needle deflection, counterclockwise for left
needle deflection.
Note
The heading trim indicator will indicate a
mistrim condition whenever the autopilot is
engaged with the airplane in a bank. It will
also indicate a mistrim condition whenever
lateral trim conditions change during auto¬
pilot controlled flight. To eliminate the re¬
quirement for trimming after engagement it
is recommended that the autopilot be engaged
with the airplane in a coordinated zero-bank
attitude.
PITCH CONTROL KNOB.
A pitch control knob (figure 4-22),located on the flight
controller, is used to trim for level flight and to control
the airplane in climbs and descents when the altitude
switch is not engaged. The pitch control knob may
also be used in coordination with the turn knob for
combined maneuvers. Rotating the pitch control knob
forward lowers the nose, and rotating the knob aft
raises the nose.
ROLL TRIM KNOB.
A roll trim knob (figure 4-22) on the flight controller
is used to center the ball on the turn and slip indi¬
cator after engagement of the autopilot. Rotate the
knob clockwise for a ball-left condition, counter¬
clockwise for a ball-right condition.
Note
It will be necessary to use the roll trim knob
only if the autopilot is engaged when the air¬
craft is flying in an uncoordinated manner.
If the autopilot is engaged during uncoordi¬
nated flight, it is usually fasrer to disengage
the autopilot, trim for coordinated flight
manually, and reengage. The autopilot will
synchronize with the new flight attitude
automatically, thus eliminating the need for
using the roll trim knob.
AUTOTRIM SWITCH AND INDICATOR.
When the autopilot is engaged (autotrim switch must
be ON), the elevator trim system is operated auto¬
matically to minimize stick force at disengagement.
The elevator trim button on the control stick is de¬
energized when the autopilot is engaged. The auto¬
trim indicator (figure 4-22), located on the flight con¬
troller panel, indicates correct operation or malfunc¬
tion of the automatic trim system. When the autotrim
system is operating properly, the trim indicator pointer
will fluctuate to either side of center. If there is mal¬
function in the system the pointer will remain con¬
stantly deflected to one side. If this condition is
noted, speed should be reduced before disengaging
the autopilot; otherwise the airplane may pitch sharp¬
ly upon disengagement, thus imposing excessive
ALTITUDE SWITCH.
A solenoid-held altitude switch (figure 4-22) located
on the function selector panel connects the altitude
control to the autopilot elevator bridge circuits. When
the switch is at ON, the autopilot will fly the airplane
accurately at the pressure altitude at which it was fly¬
ing when the switch was turned to ON- For a change
in flight altitude, the switch is turned to OFF; the
airplane flown to the new altitude and trimmed for
level flight; and then the switch is placed at ON.
The altitude switch snaps to OFF if the auropilot is
disengaged or if the ILS approach switch is moved
to ON.
4*27
Section fV
TO. 1F-89H-1
Note
The altitude switch can provide limited trim;
however, the airplane should be trimmed for
approximately level flight before placing the
altitude switch at ON. When large trim
changes are required, it is necessary to retrim
manually by means of the pitch control knob
or by disengaging the autopilot, retrimming
the aircraft, and reengaging the autopilot and
altitude control.
AUTOPILOT NORMAL OPERATION,
Ground Tests.
During engine runup, turn on and engage autopilot
and perform ground check as detailed in Section II.
Normal Engaging Procedure,
The autopilot can be engaged whenever the airplane
is flying within the autopilot engaging limits. Engage
the autopilot as follows:
1. Power switch—ON (90-second warmup required).
2. Turn knob—Detent.
3. Autotrim switch—ON.
4. Trim the airplane for coordinated zero-bank
attitude within -t50-degree pitch attitude.
5. Engaging switch—ENGAGE. Switch will hold
in ENGAGE position if proper conditions for engage¬
ment have been met; otherwise switch will spring
back when released.
6. Autotrim indicator—Check that needle of auto¬
trim indicator is oscillating either side of center.
Engaging Procedure In Turns
Or Uncoordinated Flight.
When the autopilot is engaged in turns or in uncoordi¬
nated flight, it will be necessary to trim the autopilot
as follows:
1. Center the ball on the turn and slip indicator
using the roll trim knob. Rotate the knob clockwise
for a ball left condition, counterclockwise for a ball
right condition.
2. Center the needle on the heading trim indicator
using the heading trim knob. Rotate the knob clock¬
wise for a right needle condition, counterclockwise for
a left needle condition. It is usually quicker and easier,
however, to disengage the autopilot, trim the airplane
for coordinated zero-bank attitude, and reengage the
autopilot.
Autopilot Trimming Procedure,
1. Trim the airplane manually after takeoff.
2. After engaging autopilot, check the turn and
slip indicator. If a ball left or a ball right condition
exists, center the ball by rotating the trim knob
clockwise or counterclockwise respectively. Wings will
level after this and the following steps are completed.
3. After centering the bail, check the heading trim
indicator. If the needle is deflected, return it to ap¬
proximate center by rotating the heading trim knob
in the direction indicated by the needle.
If the airplane trim condition changes during flight
on autopilot, always center the ball with the roll trim
knob before centering the heading trim indicator
needle. This procedure makes precise trimming of
the autopilot possible in one operation and should
always be used.
Straight And Level Flight.
Fly to the desired altitude, trim the airplane for ap¬
proximately level flight, and place the altitude switch
ON. The autopilot will fly the airplane at the pressure
altitude existing when the switch is placed ON. (If
the altitude switch is OFF, the autopilot will maintain
the airplane in the pitch attitude established by the
pitch knob but will not necessarily maintain level
flight.) When the turn control knob is in detent,
the gyrosyn compass system establishes a heading ref¬
erence to maintain the airplane in straight and level
flight. If a lateral mistrim condition develops (such as
would be caused by an unbalanced wing tip load) the
autopilot will maintain the airplane laterally level
and in straight flight but with heading mistrim in the
direction of the heavy wing. To compensate for this
condition, center the heading trim indicator needle
using the heading trim knob.
Maneuvering Flight.
Autopilot-controlled climbs and descents can be made
using the pitch control knob. (Altitude switch should
be OFF.) Rotate the pitch knob slowly and smoothly
to change pitch attitude. If the pitch knob is rotated
rapidly, thus calling for excessive "G’s/' the "G”
limiting clutch in the elevator servo will slip, and
the airplane will not respond. To correct this situation,
disengage the autopilot, trim the airplane to the de¬
sired attitude, and reengage the autopilot. The auto¬
pilot will maintain the new attitude until changed
by means of the pitch knob. Pull-ups from shallow
dives may be made using the pitch knob, but pull-ups
from steep dives must be made manually. Coordinated
turns can be made using the turn knob. Bank angle
(and corresponding turning rate) will be proportional
to turn knob rotation. When the turn knob is re¬
turned to detent, the airplane will return to level,
ending the turn. After a 5-second delay (which allows
the airplane to stabilize on the new heading) the
autopilot will fly the airplane on the compass head¬
ing existing at that instant. Combined maneuvers
can be made by coordinated use of the turn and pitch
knobs.
Disengaging Procedure.
The autopilot may be disengaged in three ways:
squeezing the disconnect switch on the control stick,
moving the engaging switch to OFF, or moving the
4*28
T.O. 1F-89H-1
Section IV
power switch to OFF. Squeezing the disconnect switch
or moving the engaging switch to OFF leaves the
autopilot in standby status (ready to operate as soon
as it is reengaged). Moving the power switch to OFF
turns off all electrical power to the autopilot putting
it completely out of operation. If it is left off for an
appreciable length of time, a 90-second or more
warmup will be required before the autopilot can
be used again. Normally the power switch should be
left at ON at all times during flight,
AUTOPILOT EMERGENCY OPERATION.
If the autopilot fails or functions erratically, disen¬
gage the autopilot and turn the power switch to OFF,
When the autopilot is disengaged, the sideslip sta¬
bility augmemer will resume its normal function of
directionally stabilizing the airplane.
AUTOMATIC APPROACH EQUIPMENT.
receive localizer signals, however.) The approach
switch cannot be turned to OFF manually. It snaps
to OFF only when the autopilot is disengaged.
AUTOMATIC APPROACH EQUIPMENT
OPERATION.
Refer to ILS—Autopilot-Controlled Approach, Sec¬
tion IX,
ARMAMENT*
Information on this equipment is given in T.O.
1F-89H-1A, Confidential Supplement, The following
figures, 4*23 through 4-25, are also contained in the
supplement:
Figure 4-23 GENERAL ARRANGEMENT
Figure 4-24 ARMAMENT CONTROL PANEL
Figure 4-23 ARMAMENT SELECTION TABLE
Automatic approach equipment is provided in the
autopilot system. Localizer and approach couplers en¬
able the autopilot to use signals from the localizer
and glide-slope receivers of the YHF navigation set
for reference in azimuth and elevation in autopilot
controlled approaches. (The localizer coupler is not
designed for autopilot controlled flight on omni-
range and should not be used.) The signals fed to
the localizer and glide-slope couplers are the same as
those that move the vertical and horizontal bars on
the ILS course indicator. The localizer and glide-slope
couplers supply autopilot signals to maintain the
airplane at the center of the localizer and glide beams
respectively. This equipment can be disengaged in¬
stantly by squeezing the autopilot disconnect switch
on the control stick,
LOCALIZER SWITCH.
A solenoid-held localizer switch (figure 4-22) on the
function selector panel connects the localizer coupler
to the autopilot. The switch has ON and OFF posi¬
tions. When the switch is placed at ON (after the
localizer beam has been intercepted according to stand¬
ard ILS procedures), the coupler feeds signals to the
autopilot to provide automatic bracketing and beam
following. The localizer switch can be turned off man¬
ually, or will snap to OFF automatically when the ap¬
proach switch (figure 4-22) is placed at ON, or when
the autopilot is disengaged.
APPROACH SWITCH.
A solenoid-locked approach switch (figure 4-22) on
the function selector connects the glide-slope coupler
to the autopilot. The switch has ON and OFF posi¬
tions. When the switch is placed at ON, the airplane
noses down and follows the glide beam, and the local¬
izer switch snaps to OFF. (The autopilot continues to
OPTICAL SIGHTHEAD (Ml 69)*
Information on this equipment is given in T.O.
1F-89H-1A, Confidential Supplement,
E-9 FIRE CONTROL SYSTEM.
Information on this equipment is given in T.O,
1F-89H-1A, Confidential Supplement, The following
figures, 4*26 through 4-31, are also contained in the
supplement:
Figure 4-26
Figure 4-27
Figure 4-28
Figure 4-29
Figure 4-30
Figure 4-31
RADAR CONSOLE
ANTENNA HAND CONTROL
RADAR TEST PANEL
PILOT’S AND RADAR OBSERVER’S
SCOPES
RADAR INDICATOR CONTROL
PILOT S POWER-CONTROL BOX
SINGLE-POINT FUELING SYSTEM.
All fuel tanks except the pylon tanks can be fueled
through a single high-pressure fitting located on the
lower side of the tight wing, aft of the wheel well;
To prevent overflowing and to prevent high fueling
pressure from entering the tanks, a fluid level actuated
shutoff valve in each tank closes as the tank becomes
full. (See figure 4-32.) The system is designed to use
a maximum of 55 psi during single-point operations.
This airplane cannot be defueled through the use of
the single-point fueling system.
SINGLE-POINT FUELING CONTROLS.
A 28-volt d-c tip tank control switch and a tank shut¬
off precheck switch (figure 4-33) are located in a
4-29
Section IV
T.O. 1F-89H-1
Figure 4-32,
4-30
T*0. 1F-89H-1
Section IV
switch box mounted in the right main landing gear
wheel well A 28-volt d-c circuit breaker located on
the pilot's fuel control panel protects the single-point
fueling system. The circuit breaker is left IN for all
operations. The tip tank control switch has FILL and
CLOSED positions. Placing the switch at FILL permits
the tip tanks to fill during single-point fueling; placing
the switch at CLOSE causes the tip tank shutoff valve to
close and prevents fuel from entering the tip tanks
during single-point fueling. The tank shutoff precheck
switch has CHECK and NORMAL positions and is
spring-loaded to NORMAL. Holding the switch at
CHECK causes the tank shutoff valve in each tank
to close under the same conditions that cause the
valves to close when the tanks become full* While
the precheck switch is held at CHECK, a rate of flow
of approximately 40 gpm (noted on the single-point
fueling equipment) indicates that all shutoff valves
are functioning and that the system is safe for single¬
point fueling. Failure of any tank shutoff valve to
close will be evidenced by a rate of flow in excess
of 100 gpm, and single-point fueling must be stopped
immediately. Releasing the tank shutoff precheck
switch to NORMAL causes the tank shutoff valves to
reopen and single-point fueling to continue. When the
switch box cover is closed, the tip tank and tank shutoff
precheck switches are automatically positioned for
normal fuel system operation.
CAUTION
SINGLE-POINT
FUELING PANEL
Fuel selector switches must be at ALL
TANKS or PUMPS OFF during single-point
deling. A WING TANKS selection will al¬
low fuel under high pressure to enter and
damage low-pressure engine fuel components.
SINGLE-POINT FUELING OPERATION.
1. Apply 28-volt d-c external power*
2. Connect fueling nozzle to single-point fueling
adapter.
3- Pressure refueling circuit breaker—IN,
4* Fuel selector switches—ALL TANKS if engines
are operating; PUMPS OFF if engines are shut down,
5. Tip tank control switch—FILL or CLOSED, as
desired*
6. Single-point fueling nozzle valve—OPEN; then
immediately hold precheck switch at CHECK. If rate
of flow- does not exceed approximately 40 gpm after
12 seconds, release tank shutoff precheck switch to
NORMAL and allow airplane to be fueled; if rate
of flow is 100 or more gpm after precheck switch is
held at CHECK for 12 seconds, indicating a shutoff
valve failure, stop single-point fueling immediately.
Figure 4-33,
Under no circumstances must single-point
fueling be continued if a tank shutoff valve
fails to dose during prechecking. Failure of
a shutoff valve to close during single-point
fueling will result in structural damage and
a serious fire hazard.
Note
Failure of only one tank shutoff valve 'will
result in a rate-of-flow greater than 100 gpm,
7* After airplane ts fueled, turn fueling nozzle
valve off, remove fueling nozzle, and close single¬
point fueling switch box door,
MISCELLANEOUS EQUIPMENT.
WINDSHIELD WIPER.
The windshield wiper operates on 28-volt d-c power*
The windshield wiper switch (figure 1-9) turns the
4-31
Section IV
TO. 1F-89H-1
wiper on and off, and has ON, OFF, and PARK posi¬
tions. The switch is adjacent to the speed rheostat
(figure 1*9) which is located above the pilot's left con¬
sole, The speed rheostat has INC and DEC positions
for controlling the speed of the wiper motor. The
speed rheostat must be at INC before the windshield
wiper switch is turned to ON. If the wiper blade stops
at an undesirable position when the switch is turned
to OFF, the switch can be held momentarily to the
spring-loaded PARK position; the blade will move to
the right and stop automatically. If the wiper blade
stops and cannot be started with the speed rheostat, the
wiper should be turned off.
The speed rheostat should not be used to stop
the wiper. Before either stopping or starting
the wiper, the speed rheostat should be turned
to INC.
RELIEF TUBES.
The relief tube for the pilot is on the floor to the
right of his seat; one for the radar observer is to the
right of his seat, aft of the wing spar.
MISCELLANEOUS PARTS STORAGE.
Fuselage and wing jack pads, mooring fittings, and
microphones are stored in two bags in the radio and
equipment section in the aft fuselage. The ground
safety locks and the pitot tube covers are in a third
bag near the floor to the left of the radar observer's
seat.
MAP AND DATA CASE5.
A data case and flight report holder (figure 1-9) is
beside the pilot's left console. A map data case (figure
4-8) is beside the right console in the radar observer*®
cockpit, and an airplane data case is in the aft radio
and equipment section. Two spring clips are located
on the upper right surface of the pilot's glare shield
to be used as required for temporary storage of maps,
computer, flight plan, etc. while the pilot is navigating.
CHECKLISTS.
Each crewmember has a permanently installed metal
checklist in his cockpit. The pilot's checklist (figure
1-11) slides out at the top of the center pedestal; the
radar observer’s checklist (figure 4-6) is above his
instrument panel.
REAR VIEW MIRRORS.
A mirror on the left frame of the windshield enables
the pilot to see rearward. A mirror on the right side
of the canopy frame allows the pilot and radar observer
to see each other.
EMERGENCY SIGNAL SYSTEM.
A red light and spring-loaded button-type switch in
each cockpit provide a visual emergency system for
the pilot and radar observer. In case of interphone
failure or loss of the canopy, each crewmember can
communicate with the other by means of code or
prearranged signals. In the pilot's cockpit the button
and signal light (figure 1-12) are mounted on a bracket
directly below the right canopy defog duct. In the
radar observer's cockpit the button (figure 4-8) is
mounted on the inboard side of the right console and
the signal light (figure 4-6) is located below the left
side of the main instrument panel. The system is
powered by the 28-voit d-c primary bus,
BLIND FLYING CURTAIN ASSEMBLY.
Five orange acetate curtains can be snapped onto
fasteners mounted in the pilot's cockpit to mask the
windshield and forward canopy for simulated instru¬
ment flight. The curtains will not obstruct normal
vision; but when the pilot wears blue goggles, he is
unable to see through the curtains.
ANTI “G“ SUIT EQUIPMENT.
The pilot's and radar observer’s anti "G" suits are
inflated by air pressure from the engine compressors.
An anti "G" suit intake tube attaches to an air
pressure outlet on the front of each seat. A pressure
regulator valve (figures 1-9 and 4-7) to the left of the
pilot's and radar observer's seats is moved to LO and
HI to control the pressure in the suit. Acceleration
above 1,75 "G’s” causes the valve to open, inflating
the anti ,f G" suit; for each additional "G” accelera¬
tion, the suit is inflated LO psi (LO setting) or 1.5
psi (HI setting). A button on top of the valve can
be pressed to inflate the suit momentarily.
4-32
T.O. 1F-89H-1
Section V
TABLE OF CONTENTS
Page
Minimum Crew Requirements.* * - - 5-1
Engine Limitations .*..*•*♦..* - * 5-1
Airspeed Limitations.5-7
Canopy Limitations * * . *. 5-10
Prohibited Maneuvers. 5-10
Acceleration Limitations.*.* 5-10
Center-of-Gravity Limitations.- - 5-15
Weight Limitations * *.. ..* ■ - 5-15
INTRODUCTION.
Cognizance must be taken of instrument markings shown
on figure 5-1, since they represent limitations that are
not necessarily repeated in the text.
MINIMUM CREW REQUIREMENTS.
The minimum crew is a pilot for local day VFR
flights, A radar observer or a qualified crew member
will be added for cross country, night, or IFR flights,
or at the discretion of the commander for other
operations,
ENGINE LIMITATIONS.
STARTING (AIRPLANES EQUIPPED WITH J35-35
ENGINES).
During starting, the maximum allowable exhaust gas
temperature is 915°C Exhaust gas temperatures be¬
tween 750°C and 915°C inclusive are permitted for no
more than 20 seconds. On afterburner starts, if the ex¬
haust gas temperature momentarily exceeds 915°C or
if 5 seconds after the start the exhaust gas temperature
exceeds 750°C, stop the afterburner,
STARTING (AIRPLANES EQUIPPED WITH
J35-35A ENGINES).
During starting, the maximum allowable exhaust gas
temperature is 9G0°C Exhaust gas temperatures be¬
tween 735°C and 900°C are permissible for no more
than 20 seconds. On afterburner starts, if the exhaust
gas temperature momentarily exceeds 900°C or if 5
seconds after an a fterburner start, the exhaust gas
temperature exceeds 735°C, stop the afterburner* Nor¬
mal power is 95*6% rpm; military power is 100%
rpm without afterburning; and maximum power is
100% rpm with afterburning. There are no engine
operating time limits,
ACCELERATION (AIRPLANES EQUIPPED WITH
J35-35 ENGINES).
During accelerations, the momentary exhaust gas tem¬
perature is not to exceed 915°C; but a peak tempera¬
ture between 915° and 940°C is permitted for a maxi¬
mum of 3 seconds at engine speeds below 75% rpm.
Temperatures between 750°C and 915°C inclusive are
permitted for no more than 20 seconds. The engine must
be removed for overhaul if speed momentarily exceeds
104% rpm or 103% rpm stabilized with or without
excessive exhaust gas temperature. Stabilized engine
speeds greater than 103% rpm or 104% momentary
rpm are prohibited and engine must be removed for
overhaul if these limits are exceeded. The throttle
must be reset if stabilized engine speed exceeds
102% rpm*
5-1
Section V
T.O. TF-89H-T
195 KNOTS MAXIMUM FOR FULL FLAPS OR
LANDING GEAR DOWN.
BELOW 20,OOQ-FOOT PRESSURE ALTITUDE. THE
AIRSPEED LIMITATION IS 470 KNOTS IAS OR
MACH 0,90. WHICHEVER IS LESS.
THE INSTRUMENT SETTING IS SUCH THAT THE
RED POINTER WILL MOVE TO INDICATE THE
LIMITING STRUCTURAL AIRSPEED OR THE
AIRSPEED REPRESENTING THE LIMITING MACH
NUMBER, WHICHEVER IS LESS.
msntmm markings
-2.33 G” MAX WITH EMPTY TIP AND
PYLON TANKS AND WITH MISSILES
RETRACTED ABOVE 12,000 FEET,
-1.67 'G“ MAX WITH ANY AMOUNT OF
TIP OR PYLON FUEL AND WITH MISSILES
RETRACTED ABOVE 12.000 FEET.
+ 3.67 "G" MAX WITH ANY AMOUNT OF
TIP OR PYLON FUEL AND WITH MISSILES
RETRACTED ABOVE 12.000 FEET.
+ 4.50 "G" MAX WITH EMPTY TIP AND
PYLON TANKS AND WITH MISSILES
EXTENDED ALL ALTITUDES.
+ 5.00 "G" MAX WITH EMPTY TIP AND
PYLON TANKS AND WITH MISSILES
RETRACTED BELOW 12.000 FEET.
+ 5.67 “G" MAX WITH EMPTY TIP AND
PYLON TANKS AND WITH MISSILES
RETRACTED ABOVE 12,000 FEET.
ACCELEROMETER
MACHMETER
BELOW 20.OOQ-FOOT PRESSURE
ALTITUDE THE AIRSPEED
LIMITATION IS MACH 0.90 OR
470 KNOTS IAS, WHICHEVER
IS LESS.
H-S4(1)B
5-2
Figure 5-1 (Sheet 1 of 5).
T.O. 1F-B9H-1
Section V
ENGINE TACHOMETER
49 %— 51 % IDLE LIMITS
SOb-95% OPERATING RPM RANGE
lOO^a MAXIMUM
Based on all fuel grades
OIL PRESSURE
15 PS1 MINIMUM FOR FLIGHT
25-45 PSI CONTINUOUS OPERATION
45 PSI MAXIMUM FOR FLIGHT
EXHAUST TEMPERA
J35-35A
J35-35
MINIMUM FOR FLIGHT 315 C
CONTINUOUS OPERATION 3t5-629°C
315 C MINIMUM FOR FLIGHT
31 5“6SO°C CONTINUOUS OPERATION
MAXIMUM FOR FLIGHT 733°C
MAXIMUM DURING STARTING
AND ACCELERATION ONLY
900°C
75Q q C MAXIMUM FOR FLIGHT
0 MAXIMUM DURING STARTING
AND ACCELERATION ONLY
Figure 5-1 (Sheet 2 of 5).
5-3
Section V
T.O, TF-89H-1
INSTRUMENT MARKINGS
H-«4(3)
NOSE GEAR BUNGEE
PRESSURE
720 PS] MINIMUM ■■■■
720-780 PSi OPERATING RANGE
780 PSI MAXIMUM
5-4
MAIN GEAR BUNGEE
PRESSURE
675 PSI MINIMUM
675-775 PSI OPERATING RANGE
775 PSt MAXIMUM
Figure 5-T (Sheet 3 of 5 )*
VOLTMETER AC
no-120 VOLTS
OPERATING RANGE
150 VOLTS MAXIMUM
LOADMETER
28 -VOLT DC
1.0 CONTINUOUS
OPERATING LIMIT
VOLTMETER DC
TO. 1F-89H-1
Section V
II YURA UUC RESER VOIR
PRESSURE GAGE
8-12 PSI OPERATING RANGE
PILOTS SEAT
PRESSURE GAGE
1600 PSI MINIMUM
1600—1800 PSI
OPERATING RANGE
BRAKE
ACCUMULATOR
CANOPY EJECTOR
PRESSURE
1500—2000 PSI
^ : ! OPERATING RANGE
800 PSt ONE APPLICATION
REMAINING
2500 — 3500 PSI NORMAL
3500-4100 PSI ABOVE NORMAL:
ALLOWABLE
4100 PSI MAXIMUM
LEFT HYDRAULIC
SYSTEM
1000-2500 PSI MOMENTARY
ALLOWABLE
2500-3050 PSI NORMAL
3150 PSI MAXIMUM
RIGHT HYDRAULIC
SYSTEM
1000-2500 PSI MOMENTARY ALLOWABLE
2500-3050 PSI NORMAL
3150 PSI MAXIMUM
Figure 5- J (Sheet 4 of 5J*
5-5
LOW
19-22 PStA
HIGH
28-31 PSIA
RADAR
jl nESSURE
ENGINE SCREEN
SYSTEM PRESSURE
LEFT SYSTEM
1500-1800 PSf
OPERATING RANGE
1800 PSI MAXIMUM
RIGHT SYSTEM
T 500-1800 PS I
OPERATING RANGE
Mi 1800 PS! MAXIMUM
INSTRUMENT MARKINGS
BRAKE ACCUMULATOR
AIR PRESSURE
maximum for flight
EMERGENCY
AIRBRAKE
PRESSURE
1500-1800 PSI
OPERATING RANGE
.SHER
PRESSURE
400-440 LBS,
OPERATING RANGE AT 70°F.
Figure 5-1 fSfreef 5 of 5J.
^04(5)
T.O. T F-89H-1
Section V
MISSILE LAUNCHER ACCUMULATOR
AIR CAGE
%77"H
St
PRE55 5 V C>7
TEMPERATURE: F
AIR FILLER CHART
OPERATING RANGE
11OO 1200 1300 1400 1500 1600
GAGE AIR PRESSURE: PS1 H-m
Figure 5-2.
ACCELERATION (AIRPLANES EQUIPPED WITH
135-35A ENGINES}.
The following J35-35A operating temperature limits
must be observed: During accelerations, the momen¬
tary exhaust gas temperature is not to exceed 9Q0°C,
except that peak temperatures between 900°C and
925°C are permitted for a maximum of 3 seconds at
engine speeds below 75% rpm. Temperatures between
735 °C and 90Q°C are permissible for no more than 20
seconds. The engine must be removed for overhaul if
speed momentarily exceeds 104% rpm or 103% rpm
stabilized with or without excessive exhaust gas tem¬
perature, Engine speeds greater than 103% rpm are
prohibited and engine must be removed for overhaul
if this rpm is exceeded under stabilized conditions or
104% rpm is momentarily exceeded. Have the throt¬
tle reset if stabilized engine speed exceeds 102% rpm,
IXHAUST GAS TEMPERATURE VERSUS AMBIENT
TEMPERATURE.
Abnormally low exhaust gas temperatures for the
existing ambient temperature will result in a loss of
thrust. Available thrust may be insufficient for take¬
off under this condition on a runway of limited length.
Refer to figure 5-3 to ensure that exhaust gas tern-
features and runway temperature are within limits
which allow sufficient thrust for takeoff.
Note
Ambient temperature does not effect peak
exhaust gas temperature limits.
ALTERNATE HPEEL LIMITATIONS.
If MIL-F-5572 aviation gasoline is used as an alternate
fuel, the following limitations must be observed:
L With ambient temperatures of 0°F ( — 18°G) and
lower, do not exceed Mach 0,4 below 5000 feet with af¬
terburners operating,
2, With sea level ambient temperatures exceeding
70° F (21°C), do not exceed 25,000-foot altitude.
These limitations are to prevent cavitation of the engine-
driven and booster pumps,
AIRSPEED LIMITATIONS*
Pending completion of static and flight tests, the
airspeed limitations are as follows:
1, Below 20,000-foot pressure altitude, airspeed is
restricted to 470 knots IAS or Mach 0,90, whichever
is the lower indication. These limits are imposed to
prevent excessive structural loads resulting from gusts.
Above 20,000-foot pressure altitude, airspeed is un¬
restricted,
2. The preceding restrictions apply to all fuel and
armament loading conditions with the following excep¬
tion: with any amount of usable tip tank fuel, less than
a full load of rockets (or approved dummy), and less
than a full load of missiles do not exceed 400 knots
indicated airspeed at any altitude. If a full comple¬
ment of either type armament is aboard the airplane,
the 400-knot indicated airspeed restriction does not
apply.
Note
Cruising at 400 knots indicated airspeed in¬
stead of the airspeeds recommended in the
Flight Operation Instruction Charts has neg¬
ligible effect on range.
5-7
Section V
T.O. 1F-89H-1
EXHAUST GAS TEMPERATURES
VS
AMBIENT TEMPERATURES
Air in lei screens extended
Without afterburning
J35-35 ENGINES
NOTE: Afterburning lowers exhaust
gas temperatures up to 5°C.
J35-35A ENGINES
EXHAUST GAS TEMP
100% RPM
°C MAX °C MIN
AMBIENT TEMP
°C °F
EXHAUST GAS TEMP
100% RPM
°C MAX °C MIN
749
729
38
IOO
735
715
749
729
32
90
735
715
743
723
27
so
729
709
736
716
21
70
721
701
72S
708
16
60
713
693
719
699
IO
50
705
685
711
691
4
40
697
677
702
682
30
687
667
692
672
-7
20
678
658
683
663
-12
10
669
649
673
653
“18
O
659
639
66 3
643
-23
-10
649
629
653
633
-29
-20
639
619
645
625
-34
-30
632
612
633
613
-40
-40
623
603
631
6! 1
-46
-50
616
596
626
606
-51
-60
611
591
98%
RPM
98% RPM
748
728
43
IIO
733
713
749
729
49
120
735
715
H-l
5-8
Figure 5-3.
T.O. 1F-89H-1
Section V
PSI 80 90 100 110 120 130 140 ISO 160 170 180 190 200 210 220 230
MAIN LANDING GEAR TIRE PRESSURE VARIATION + 10 PSI
NOTE
NOSE LANDING GEAR
TIRE PRESSURES:
160 PSI TO 38,000 POUNDS
190 PSI ABOVE 38,000 POUNDS
r AIRPLANE "
GROSS WEIGHT
. IN POUNDS ^
H-l 1 7S
Figure 5-4.
3* With less than full armament and less than 400
pounds of fuel remaining in the main tanks, do not
exceed 400 knots indicated airspeed.
AUTOPILOT IMITATIONS,
Au topi lot-con trolled flight below 25,000 feet pressure
altitude is limited to 425 knots IAS or Mach 0.78,
whichever is lower; however, the 400-knot IAS limi¬
tation described in the Airspeed Limitations paragraph
must be observed.
WING FLAP LIMITATIONS,
Do not exceed the following structural limit airspeed
of the wing flaps, or the wing flaps may fail struc¬
turally:
Wing Flap Positions IAS—Knots
Wing flaps at takeoff {gear up) 230
Wing flaps full down (gear up or down) 195
Note
A wing flaps full down and 195 knots IAS
condition can occur only when the airplane
is accelerated to 195 knots IAS after extending
the flaps. Airloads prevent fully extending
the flaps at or above this airspeed.
LANDING GEAR LIMITATION,
With the wing flaps in any position, the structural
limit airspeed of the landing gear and main landing
gear doors is 195 knots IAS and 1.2 "GY* during
retraction.
TIRE LIMITATION,
Speed on the ground should not exceed 140 knots at
takeoff or 122 knots at landing to obtain normal tire
life. Exceeding these speeds on occasion will not
necessarily result in tire failure; however, continual
operation at excessive ground roll speeds will result
in reducing tire life and premature failure. For tire pres¬
sures see figure 5-4*
LANDING—TAXI (LIGHT LIMITATION,
Do not extend landing light above 175 knots IAS*
The light was designed for use only during final
approach and landing. If this limitation is exceeded,
the light may fail structurally.
PYLON LIMITATIONS,
Overlimit stresses on the wing pylon racks may occur
in flight: if a tank collapses; if the acceleration limit
for the airplane with pylon fuel (1000 pounds or more)
is exceeded; or if the airplane exceeds a roll rate of 90
degrees per second with 1000 pounds or more of pylon
5*9
Section V
TO, 1F-89H-1
fuel. Landing with pylon tank fuel is prohibited.
Rough taxiing with pylon fuel aboard should be noted
on DD Form 781,
PYLON TANK JETTISON LIMltATIONS.
The following airspeed limitations will apply when
jettisoning pylon tanks:
1. When using power ejection procedure, do not
exceed 300 knots IAS.
2, When releasing tanks to fall by gravity, fly
the airplane as near as possible to 200 knots IAS,
CAUTION jj
# Releasing empty tanks at speeds substantially
above or below 200 knots IAS, or ejecting
them at airspeeds above 300 knots IAS may
cause the tanks to tumble and strike the air¬
plane.
# To prevent damage to the speed brakes, it
is recommended that they be closed when
jettisoning the pylon tanks*
# When the pylon tanks are jettisoned manually
(gravity drop), minor damage to the airplane
may result,
CANOPY LIMITATIONS*
Speeds must not exceed 50 knots IAS when airplane
is taxied with the canopy open,
PROHIBITED MANEUVERS*
SPINS.
Intentional spins, with or without external stores,
are prohibited,
ACROBATICS.
Acrobatics will not be performed below 12,000 feet,
INVERTED FLIGHT.
Inverted flight can be maintained without afterburn¬
ing for approximately S seconds at 20,000-foot pressure
altitude, because of the limited amount of fuel avail¬
able to the engines. At the time the airplane is inverted
only that fuel already in the fuel lines, fuel pumps and
fuel controller will be available for use; when that has
been used flameout will occur. At lower altitudes this
time will be considerably reduced because of increased
fuel consumption.
Note
Inverted flight or any other maneuver involv¬
ing negative ^G” forces with maximum power
will result in immediate afterburner flameout,
LANDING.
Landing with any tip tank or pylon tank fuel is prohib¬
ited. Landings at heavier than normal landing weight
should be made with caution. Normal landing weight is
one half or less of internal fuel and no tip pod arma¬
ment, These limitations are imposed to avoid overstress¬
ing the pod attachment fittings,
ROCKET/MISSILE FIRING.
Firing rockets or missiles with any amount of tip fuel
is prohibited* This limitation is imposed to prevent
overloading tip pod attachment fittings,
AILERON AND RUDDER MOVEMENT.
The following restrictions to aileron and rudder
movement apply except during takeoff and landing:
1, With any pylon or tip tank fuel, other than
residual, do not exceed one-third full aileron deflection.
Note
With no tip tank fuel and no pylon fuel (with
or without empty pylon tanks), aileron de¬
flection is unrestricted.
2. When pylon tanks (empty or full) are carried,
or when any tip tank fuel is carried, abrupt rudder
deflections are prohibited.
Note
Without pylon tanks or tip tank fuel, rudder
deflection is unrestricted, except for fish-
tailing maneuvers,
ACCELERATION LIMITATIONS.
A load factor envelope, shown on the Operating Flight
Strength Diagram (figure 5-5), includes the operating
gross weight and operating altitude ranges of the
airplane. Lines on the left of the charts represent
maximum lift limitations; top and bottom lines specify
structural limit load factor; lines on the right indicate
limit airspeeds or elevator control boundaries. The
elevator control boundary lines show the necessity
for careful regulation of airspeed during dive maneu¬
vers because a small increase in IAS will result in a
noticeable decrease in available load factor or ability
to maneuver* This effect will be dangerous as speeds
increase above the maximum level flight airspeed.
5-10
T.O. 1F-89H-1
Section V
SEA LEVEL i
12,000 FEET>
20,000 FEETi
30,000 FEET<
40,000 FEET 1
How to use charts
SOLID LINES REPRESENT STALL LIMIT
BROKEN LINES REPRESENT ELEVATOR
CONTROL POWER LIMITS
Select an indicated airspeed.
Move up the chart to a selected altitude
(solid or broken line}.
Move to the left to find the maximum
number of “CV* you can pull
at that airspeed and altitude. H-8BHJC
FOR SYMMETRICAL FLIGHT IN SMOOTH AIR
STRENGTH DIAGRAM
NO INTERNAL FUEL
NO TIP TANK FUEL
NO PYLON TANKS
FULL TIP POD ARMAMENT
* (Approximately pounds with empty pylon tanks)
APPROXIMATELY
POUNDS GROSS WEIGHT*
14 August 1937
Flight test
5.0 “G” LOAD FACTOR
MISSILES RETRACTED
BELOW 12,000 FT a
*5.67 “C” LOAD FA
-MISSILES RETRACJ
ABOVE 12,000 FT
j 500
- 2.00 "-G” LOAD FACTOR
MISSILES RETRACTED
BELOW 12.000 FT
OPERATING EUGRT
DATA AS OF:
DATA BASIS:
400
- 2,33 -G” WAD FACTORi
MISSILES RETRACTED
Figure 5-5 (Sheet 7 of 3j,
5-11
Section V
T.O. 1F-89H-1
FOR SYMMETRICAL FLIGHT IN SMOOTH AIR
FULL INTERNAL FUEL
NO TIP TANK FUEL
NO PYLON TANKS
FULL TIP POD ARMAMENT
APPROXIMATELY
POUNDS GROSS WEIGHT '
DATA AS OF: 14 August 1957
DATA BASIS: Flight test
STRiHGTH DIAGRAM
How to use charts,
SOLID LINES REPRESENT STALL LIMIT
BROKEN LINES REPRESENT ELEVATOR
CONTROL POWER LIMITS
Select an indicated airspeed.
Move up the chart to a selected altitude
(solid or broken line),
Move to the left to find the maximum
number of U GV 7 you can pull
at that airspeed and altitude.
Figure 5-5 (Sheet 2 of 31
5-12
T.G. IF-89H-1
Section V
FOR SYMMETRICAL FLIGHT IN SMOOTH AIR
FULL INTERNAL FUEL
FULL TIP TANK FUEL
FULL PYLON TANK FUEL
FULL TIP POD ARMAMENT
APPROXIMATELY
POUNDS GROSS WEIGHT
DATA AS OF: 14 August 1957
DATA BASIS: Flight test
0 TOO 200 300 400 500
OP MATING PLIGHT
STRENGTH DIAGRAM
How to use charts.
SOLID LINES REPRESENT STALL LIMIT
BROKEN LINES REPRESENT ELEVATOR
CONTROL POWER LIMITS
Select an indicated airspeed.
Move up the chart to a selected altitude
(solid or broken line).
Move to the left to find the maximum
number of *‘GV* you can puli
at that airspeed anil altitude. H-8K3} A
Figure 5-5 (Sheet 3 of 3 )>
5-13
Section V
T.O. 1F-89H-1
earn wtKHTs
K-'
^ ‘vl.j iiV 'x * J /Sj 1 tr>
Full internal fuel,
rockets, and missiles.
Full Lip tank
fuel added.
Pylons, pylon tanks,
and full pylon fuel added.
43.200 4Z40&
figure 5 - 4 ,
l. With any amount of tip pod armament, no tip
tank fuel, and no pylon tank fuel, do not exceed the
following load factors:
MISSILES RETRACTED
Symmetrical maneuvering flight
(above 12,000 ft) + 5,67 or -233 *GV*
Symmetrical maneuvering flight
(below 12,000 ft) + 5.00 or -2.00 ’W
Asymmetrical maneuvering flight
(above 12,000 ft) + 3-40 or 0 "GY'
Asymmetrical maneuvering flight
(below 12,000 ft) + 333 or 0 "GY'
MISSILES EXTENDED
All Altitudes
Symmetrical maneuvering
flight +430 or -L67"GY’
Asymmetrical maneuvering
flight +3-00 or 0 "GY 1
+3-00 or 0 "GY 5
2, With any amount of tip pod armament, full in¬
ternal fuel, any amount of tip tank fuel, and any
amount of pylon tank fuel, do not exceed the following
load factors:
MISSILES RETRACTED
Symmetrical maneuvering flight
(above 12,000 ft) + 3-67 or -1.67 “GY*
Symmetrical maneuvering flight
(below 12,000 ft) +3*67 or -L67 "GY 1
Asymmetrical maneuvering flight
(above 12,000 ft) +2.45 or 0 "GY’
Asymmetrical maneuvering flight
(below 12,000 ft) +2.45 or 0 "GY*
Asymmetrical maneuvers are those which
create unequal airloads resulting from aileron
or rudder deflection. A coordinated turn, how¬
ever, is a symmetrical maneuver once bank
angle is established.
Above the maximum level flight airspeed, the
maximum allowable negative load factor re¬
duces as airspeed increases, reaching —1.00
”G“ at maximum airspeed attainable.
WARNING
If airplane is trimmed for high speed flight at
low altitude, airplane will nose down sharply
if speed is reduced.
It is possible to overstress the tip tank attachment fit¬
tings and the pylon racks if a landing is made with fuel
in these tanks; therefore, tip tanks must be emptied and
5-14
T.O. 1F-89H-I
Section V
or less, and 24 percent MAC at 48,000 pounds gross
weight, varying linearly between these points. The
normal operating aft limit is 25-8 percent MAC at
48,000 pounds gross weight and 27,7 percent MAC
at 29,000 pounds gross weight; this limit varies linearly
with gross weight. It is allowable for the aft eg limit
to move 0,70 percent MAC aft of its normal position
provided that a full load of fuel is carried (with or
without pylon fuel tanks), and no tip pod rockets, ap¬
proved dummy rockets, nor missiles are carried. For
detailed instructions of weight and balance refer to
T.O. 1-1B-40 and TO. 1F-89H 5.
WEIGHT LIMITATIONS,
The forward eg limit is at 20 percent of the Mean There are no weight limitations. See figure 5-6 for
Aerodynamic Chord at 35,000 pounds gross weight design, alternate, and maximum gross weights.
pylon tanks must be emptied or jettisoned before
landing. Not all of the tip tank fuel can be dumped
during dives or deceleration because the fuel will shift
and uncover the dump tube before the tank is
emptied.
Because of the fire hazard, do not fire arma¬
ment while tip tank fuel is being dumped.
CENTER-OF-GRAVITY LIMITATIONS*
5-T5
T.O, 1F-89H-1
Section VI
SECTION VI
rum CHMMmmm
HF4B
FAB5LE OF CONTENTS
Page
Introduction , * . ....*.. ■ ■
Stolls ..- - ■ ■ 6-1
Spins . . , .....
Flight Controls.* - - -. 6-2
Level Flight Characteristics , *.* * ■ ■ ■ 6-5
Maneuvering Flight ...... .* 6-6
Diving 6-7
Flight with Asymmetrical Loading .. 6-16
Flight with External Loads.6-16
INTRODUCTION.
The airplane is a large, high speed, fast-climbing all-
weather interceptor. The two-engine design increases
dependability and permits high performance while
carrying the heavy load of armament and equip¬
ment necessary for an intercept mission. All flight
control surfaces are 100 percent hydraulically actuated.
Full-powered controls permit accurate control of the
airplane at airspeeds which would otherwise make
control forces prohibitively high. They also prevent
sudden airload changes on control surfaces from affect¬
ing the stick or rudder pedals. The wide range of speed
control possible with split-aileron speed brakes in¬
creases combat effectiveness. The sideslip stability aug-
menter provides satisfactory damping of the high speed
Dutch Roll, assists the pilot in making coordinated
turns in combat maneuvers, and provides a stable firing
platform at high speeds. Tip pod fins* in addition to
decreasing wing twist and keeping the center of span-
wise lift more nearly constant, add to the longitudinal
stability and control characteristics of the airplane. The
fins increase the stick force per "G”, particularly for the
aft eg conditions in the airspeed range where maneuver¬
ing stability is critical (from approximately 0.70 to 0.80
Mach number). Power response to throttle adjustment
is slow, as in all jet airplanes because of the high inertia
of the engine rotors. However, rapid changes of effec¬
tive power are obtainable by stabilizing airspeed at a
power setting higher than required by use of partially
opened speed brakes, then quickly changing speed
brake position as changes in effective power are re¬
quired, Excess power is greatest at medium to high
airspeeds. Consequently, to perform any maneuver
involving altitude and airspeed changes, maintain me¬
dium to high airspeeds.
STALLS.
The stall in this airplane is a mild pitch down, with
drop off usually to the left. See figure 6-2 for stall speeds
for clean landing and takeoff configurations. At low
altitudes, power-on stall IAS is approximately 3 knots
lower than power-off stall IAS for the configurations
indicated in the Stall Speed Chart, The airspeeds shown
in the chart for the landing and takeoff configurations
are for idle power. Ailerons and rudder retain sufficient
effectiveness to maintain adequate control during a
stall. Recovery from a stall is made by lowering the
nose slightly and adding power as may be required.
The altitude lost in a stall will be approximately 500
feet. Landing gear position does not affect stall speed.
6-1
Section VI
T.O. 1F-89H-T
MACH
mmm chart
50 45 40 35 30 25 20 15 10 5 0
\\\\ \ \ \ \ 1 \ l
DATA AS OF: 14 August 1957
DATA BASIS: Flight Test H- 92 B
Figure 6-1 .
Speed brake position affects stall speed as follows:
with wing flaps up, stall IAS decreases as speed brake
opening increases, reaching maximum decrease of 6
knots with speed brakes fully open. With wing flaps in
the landing position, no change in stall IAS occurs
until speed brakes are 30 degrees open; then stall IAS
increases as speed brake opening increases, reaching a
maximum increase of 7 knots with speed brakes fully
open.
ACCELERATED STALLS.
At airspeeds above Mach 0.25 the accelerated stall
region {shown by the sloping lines on the left of the
Operating Flight Strength Diagram, figure 5-5) is
characterized by buffeting, pitching, and rolling, which
increase as load factor increases. Any increase of load
factor after buffet onset is accompanied by rapid loss of
airspeed and extreme buffet. For this reason, the buffet
region should not be penetrated beyond a mild buffet.
It is recommended that accelerated stalls be practiced so
that they may be anticipated by feel of the airplane,
SPINS,
Intentional spins are prohibited. Damage to the air¬
plane's heavy complement of electronic equipment
may occur from the unusual loads developed in spins.
The airplane will not spin inadvertently and has no
dangerous inherent spin characteristics. However, be¬
cause of the air plane* s high wing loading, consider¬
able altitude will be lost during a spin. Total altitude
lost during spins varies from about 3000 feet between
stall and complete recovery for a one-turn power-off
spin in landing configuration, to about 12,000 feet for
a three-turn spin with continuous power in clean con¬
figuration. A three-turn power-off spin in clean con¬
figuration generally requires about 10,000 feet total
altitude. With the use of conventional spin recovery
technique, recovery characteristics are normal. Re¬
covery from a three-turn power-off spin in clean con¬
figuration requires between one-half and three-quarter
turn, and recovery from a one-turn power-off spin in
landing configuration requires from one-quarter to
one-half turn. With power on, the rate of recovery is
slightly slower. The conventional spin recovery tech¬
nique of full opposite rudder followed by forward stick
is normal and will produce satisfactory results; how¬
ever, a faster recovery can be effected by neutralizing
the stick at the same time opposite rudder is applied.
This method also lessens the chance of inadvertently
entering a secondary inverted spin while recovering
from a normal spin. Aileron position during the spin,
whether with the spin, neutral, or against the spin, has
no effect on the recovery. Direction of spin has no
pronounced influence on spin recovery characteristics.
Raising flaps and closing speed brakes aid spin recovery.
FLIGHT CONTROLS*
The full-powered irreversible flight control system
gives the airplane good handling characteristics. Arti¬
ficial stick feel provides a definite sense of control and
is adequate under normal conditions. Control forces
remain within moderate limits through a wide range
of airspeeds.
ELEVATOR*
Elevator control is satisfactory under normal operating
conditions. However, between Mach 0,72 and 0,78 the
elevator becomes extremely effective, and very small
deflection is required to obtain an additional "G” of
acceleration. Since the maximum power climb sched¬
ules are at these Mach numbers, more than normal
effort may be necessary in turbulent air to hold to a
close climb schedule. An elevator reversal occurs at
Mach 0.80 to 0.83 and is characterized by slight nose
6-2
T.O. 1F-89H-1
Section VI
SP££D CHART
With or without pylon tanks
GEAR UP OR DOWN
STALLING SPEED IAS KNOTS
WING FLAP
POSITION
ANGLE OF
BANK
r— POUNDS
POSITION ALTITUDE-FEET
CLOSED
CLOSED
0 9
0
117
124
132
139
147 !
5,000
ns
126
133
140
149
10 MOO
119
127
134
141
150 j
20,000
121
129
138
147
--- 1
159
2
30,000
127
133
150
161
176 ;
40 M00
140
153
165
175
187 J
45.000
149
160
169
178
189 ;
30’
0
124
132
140
147
*
157 \
45*
TO
5M00
0
TO
5 M00
136
145
153
164
175 j
60*
0°
162
175
188
201
220 ]
105
113
121
128
137 j
50°
113
T20
128
136
146
45*
125
133
141
150
162
60°
150
162
171
130
191
Vi OPEN
0 Q
100
107
114
121
129
30°
106
T14
122
130
140 i
45”
119
128
136
146
157 j
60 0
146
159
170
179
190 5j
DATA AS OF: 14 August 1957
DATA BASIS: Flight test
figure 6-2 ,
heaviness. This nose-down tendency can be trimmed
out; however, if during a turn or other maneuver, the
airspeed drops from 3 to 5 knots, the airplane will
pitch up rather sharply. At high indicated airspeeds
or at high Mach numbers, elevator control will be
limited as shown on the Operating Flight Strength
Diagram (figure 5-5). Under these conditions twisting
and bending of the airplane structure, together with
high Mach effects, cause elevator effectiveness to de¬
crease rapidly, approaching zero at sea level at approxi¬
mately Mach 0.925 (which is above the maximum
airspeed restriction of the airplane). This is due to
high dynamic pressures associated with high airspeeds
at low altitude, and high Mach number effects at
high altitude. The result is that the maximum load
factor attainable at high airspeed at a given altitude
will decrease as airspeed increases above about Mach
0.82. This means that the higher the airspeed, the
fewer the available "G’s.” At speeds of Mach 0.86 and
above, elevator effectiveness is so decreased that less
than 2 degrees of elevator deflection are available with
full stick deflection and less than 2 "GY* are available
at Mach 0.98 at 35,000 feet (an important point to
remember during a high Mach dive recovery).
If airplane control should become sluggish at
altitudes above 30,000 feet, check the hy¬
draulic reservoir pressure. If pressure is below
operating limits, reduce altitude until control
response is again normal.
6-3
Section VI
T.O. 1F-S9H-T
e, G n OVERSHOOT.
As positive or negative load factor develops on the
airplane, an elevator force-feel bob weight tends to
move the stick in the opposite direction opposing
further stick application. For each "G” increase, the
bob weight increases force against the stick 4.5 pounds.
It must be remembered, however, that if the stick is
moved abruptly, it is possible to obtain elevator posi¬
tion corresponding to high "G’s” before the "GV'
have built up on the airplane and have increased the
stick force through the action of the bobweighr. This
is apparent particularly between Mach 0.65 and Mach
0.80. Once the "G” load starts to develop, the buildup
to the point of failure can occur before corrective
action becomes effective. Thus, by abruptly pulling
back on the stick indiscriminately, it is possible to over¬
shoot the "G” limit and pull the airplane apart. When
you're at low altitudes } do not attempt abrupt pull-ups.
Do not rely upon the "feel” of the stick to keep you
out of trouble,
AILERONS.
Aileron effectiveness is adequate under all conditions
except in spins and at airspeeds above Mach 0.86 where
aileron effectiveness decreases rapidly. At an indicated
Mach number of 0.86, a slight aileron reversal occurs
which may be compensated for by using ailerons in a
direction opposite to normal. Sufficient lateral control
for performing normal maneuvers at airspeeds above
Mach 0,86 can be maintained with speed brakes opened
approximately 7-1/2 degrees. Partially opening the
speed brakes (from 10 to 20 degrees) also improves
aileron effectiveness at medium airspeeds (above Mach
0,75). At low airspeeds near the ground (such as those
used for takeoffs and landings), aileron response may
be lower than normal, particularly in turbulent air.
This condition may exist at any airplane gross weight
but can be minimized by strict adherence to nose wheel
liftoff, takeoff, approach, and landing airspeeds.
RUDDER.
Rudder operation is satisfactory under all operating
conditions. The slideslip stability augmemer should be
turned on before takeoff and left on for the duration
of the flight. This system operates automatically to
damp out any sideslipping or rolling tendencies in¬
duced by high speed and altitude effects; also, through
a signal derived from movement of the aileron con¬
trols, the system applies rudder in a turn in proportion
to aileron deflection, thereby enabling the pilot to make
coordinated turns with ailerons alone,
SPEED BRAKES.
The split-aileron speed brakes provide a much larger
drag surface than other types, making them highly
effective under all operating conditions. Lateral control
is improved at Mach numbers near cruise and above
by sightly opened speed brakes. Since the speed brakes
are symmetrical and are located almost in line with the
airplane center of gravity, their use has little effect on
trim. There is ample and positive control about all
axes with speed brakes in any position. Pitch and yaw
characteristics are not directly affected by their use.
Letdowns up to 30,000 feet per minute can be made
without exceeding 350 knots IAS, Altitude loss is
reduced by using speed brakes in high speed dive re¬
coveries; however, as the speed increases above Mach
0,90, speed brake effectiveness decreases. Above ap¬
proximately 260 knots the speed brakes will not open
fully. At Mach 0.90 they will open approximately 30
degrees only, and because of adverse compressibility
effects, little drag may result from their use. Speed
brakes are especially effective in controlling airspeed
and altitude during approach. During landing, this air¬
speed control permits fast acceleration for go-arounds.
Ground roll is reduced appreciably by moving the
speed brakes to full open after touchdown. They give
excellent airspeed control at constant throttle settings,
thus permitting high rate of closure in combat while
retaining maximum power for a fast breakaway. At
high indicated airspeeds, sufficient lateral control for
maneuvering can be maintained with speed brakes 5 to
6-4
T.O. 1F-89H-1
Section VI
tO degrees open without affecting airspeed, A 5-degree
speed brake opening will also eliminate the natural
rolloff tendency at high Mach numbers*
Note
By moving speed brake lever to the full open
position and reducing power, the airplane can
be decelerated in level flight from maximum
level flight speed to stalling speed in less
than 1 minute at any altitude*
TRIM.
Longitudinal trim is not affected by lowering the
landing gear during approach or by changes in thrust
at high airspeed* However, when shutting down after¬
burners between approximately Mach 0*84 and Mach
0*88, the high speed can no longer be maintained
(in level flight) and a push force on the stick is re¬
quired as airspeed decreases, requiring retrimming
at the lower airspeed* Nominal change in longitudinal
trim is required when changes in thrust are made at
low airspeeds* When speed brakes are opened, no imme¬
diate change in trim is required; however, as airspeed is
reduced, longitudinal trim may be necessary* The
aileron trim motor is independent of stick position.
When trimming the elevator, the trim mechanism will
not operate after the stick force is reduced to zero for
any given stick position* Elevator trim will appear
more sensitive at cruise speeds as less elevator is re¬
quired to trim for a small change in speed in this
region. Normal available rudder trim is 5 degrees left
or right. Under normal flight conditions, the emer¬
gency rudder trim knob should not be used, as the
sideslip stability augmenter system will be adversely
affected.
HIGH AIRSPEED OVERTRIM.
Stick forces vary with airspeed changes (see figure
6-3) and can be trimmed out for level flight* How¬
ever, for flight at relatively low altitudes, extreme
caution should be used in trimming out all the stick
force* If all the push force required for level flight
at relatively high airspeeds is trimmed out, and the air¬
plane then slows down, it is possible for the pull force
required for level flight (at the lower airspeed) to build
up in magnitude faster than the pilot anticipates,
causing the airplane to nose down sharply (an unsafe
attitude with the airplane close to the ground).
Do not trim out all stick force during low-level
flight at high airspeeds as the airplane may
dive sharply as airspeed is reduced.
LEVEL FLIGHT CHARACTERISTICS.
At any operating altitude and at all airspeeds, except
the range between Mach 0*80 and Mach 0*86, a push
force on the stick is required as airspeed is increased if
1 "G” flight is to be maintained. As airspeed is increased
from Mach 0.80 to Mach 0,86, 1 flight can be main¬
tained with less push force.
LOW SPEED.
The handling characteristics of the airplane at low air¬
speeds are good, except that near 1 stall, rolling
response to aileron motion may be lower than normal*
Adhere closely to nose wheel liftoff, takeoff,
approach, and landing airspeeds, especially in
turbulence or crosswinds, to assure adequate
lateral control.
CRUISING AND HIGH SPEED.
With the exception of the elevator stick force and posi¬
tion characteristics previously explained, no unusual
characteristics will be experienced in the medium to
high airspeed range. Figure 6-3 shows a typical vari¬
ation of stick force with the airplane trimmed to fly
6-5
Section VI
TO, 1F-89H-1
so
70
60
50
40
30
30
10
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y m
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*=
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—
,3
,7 .8
Figure 6-3 (Sheet T oi l).
"hands off" at cruise airspeed, and indicates the air¬
speed range of the mild reversal in normal stick force
variation.
Buffet—1 “G” Flight.
During 1 flight you will experience a mild com¬
pressibility buffet in the airspeed range from Mach
0,85 to Mach 0,90, This buffeting effect, which can be
likened to driving a car along a washboard road, is not
considered objectionable. The intensity of buffeting in¬
creases slightly with airspeed while in the buffet range,
but practically disappears above Mach 0,90.
High Airspeed Wing Drop,
At airspeeds between Mach 0,85 and Mach 0.90 (the
same range in which light buffeting is experienced in
level flight) wing drop, common to many jet airplanes
at high Mach numbers, is most likely to occur. Wing
drop may be either to the right or left, but is usually
to the left and can be eliminated by opening the speed
brakes approximately 5 degrees,
MANEUVERING FLIGHT*
STICK FORCES.
In level flight, minimum stick forces per "G” will
occur at airspeeds in the region of Mach 0,78
(see Stick Forces Chart, figure 6-3)- Because of light
stick forces^ care must be exercised when maneuvering
near this airspeed not to exceed the allowable load
factor by overcontrol. If the airplane enters accelerated
flight above Mach 0,80, the stick force necessary to pull
load factor will be high, but may be partially trimmed
out to a comfortable value. However, never trim out all
of the stick force while in accelerated maneuvers. If
enough stick force is applied and held, either by trim
or pilot effort, to pull the desired load factor, the ap¬
plied stick force will result in a rapid increase in load
factor as airspeed drops. This can result in rapidly ex¬
ceeding the design or even the ultimate load factor.
Use no more elevator trim than necessary dur¬
ing maneuvers. Use extreme caution to avoid
excessive “GY* as airspeed decreases during
high speed maneuvers.
6-6
T.O, 1F-89H-1
Section VI
LOAD FACTORS.
The maximum permissible load factor of 5.67 is the
highest allowable under any flight conditions. Above
approximately 20,000 feet it is impossible to attain 5.67
load factor because the airplane will either be forced
into an accelerated stall or the elevator control power
limit will be reached. At these altitudes, the airplane is
controllable at high Mach numbers and its flight char¬
acteristics are normal for a high performance airplane.
At medium to high airspeeds at low altitudes, the air¬
plane can be overstressed to the point of structural fail¬
ure. Because of the possibility of excessive gust loads at
low altitudes, the airplane is limited to a maximum load
factor of 5.0 below 12,000 feet. Flying at high indicated
airspeeds at low altitudes is dangerous because elevator
effectiveness, or ability to develop load factor, can
change within wide limits with relatively small changes
in airspeed. Do not attempt abrupt pullups at low alti¬
tudes, and do not rely entirely on stick feel to keep you
out of trouble. Be aware of the definite distinction be¬
tween the structural strength of an interceptor and of a
fighter-type designed for fighter versus fighter combat.
DIVING,
At any gross weight, the altitude lost during recovery
is dependent on the altitude at which recovery is
started, the angle from which the recovery is made,
airspeed during recovery, and the load factor ("GY')
held during recovery. See figure 6-4 for examples of
typical dive recovery flight paths.
Mote
Altitude lost during dive recovery as shown
in the Typical Dive Recovery illustration (fig¬
ure 6-4) and Dive Recovery Charts (figure
6-5) does not include the altitude lost enter¬
ing the dive. Dive recovery charts are based on
a constant airspeed being held during entire
recovery.
The Dive Recovery Charts (figure 6-5) show the inter¬
relation between these variables. The charts should be
studied collectively in order to understand the capabili¬
ties of the airplane and to be able to exercise proper
6-7
Section VI
T.O. 1F-89H-1
ryp/CAi om ntcovtuy
RECOVERY STARTED AT 10,000 FEET ALTITUDE AND 350 KNOTS IAS
judgment in planning dive maneuvers. The limiting air¬
speed lines on these charts represent the maximum and
minimum operating airspeeds at which the airplane
may be flown at a specific pressure altitude and for
which the load factor designated on the chart is attain¬
able. At minimum airspeeds (maximum lift lines) an
accelerated stall will occur. At airspeeds greater than
the maximum (elevator power limit lines), elevator
control is limited by aeroelastic distortion of the
airplane structure and by elevator control power
to such an extent that the airplane can no longer
develop the load factor shown on the chart. The
resultant effect causes the maximum attainable
load factor to decrease rapidly (and therefore
increases the altitude lost during recovery) for a
relatively small increase in IAS above the limiting
value shown on the chart. See figure 6-5, sheet 1 of 6
sheets, for instructions on chart use.
The altitude and IAS at which a maximum
(allowable or attainable) load factor recovery
is started should be anticipated so as not to
exceed airspeed restrictions (425 knots IAS
or Mach 0.90, whichever is the lower) and to
insure at least the minimum ground clearance.
HIGH MACH DIVE.
Performing a high Mach dive at high altitude is the
best way to become familiar with the high Mach char¬
acteristics of the airplane. This maneuver is useful in
combat for a breakaway, as an evasive maneuver, or as
an effective way to let down rapidly. Since the purpose
6-8
TO. 1F-89H-1
Section VI
of the high Mach dive is to lose altitude as rapidly as
possible, enter the dive with maximum power and at
high IAS and get into a 60-degree dive as soon as possi¬
ble-
WARNING
Generally, the steeper you dive the greater
the airspeed; however, if the angle of the dive
is steepened beyond 60 degrees, the in¬
crease in speed is negligible. Dive angles
steeper than 60 degrees result in far greater
altitude loss during recovery, A vertical dive
requires twice the altitude for recovery that
a 60-degree dive requires. At speeds associated
with high Mach dives {Mach 0,90 and above),
elevator and speed brake effectiveness are
greatly reduced. Because of the reduced elevator
effectiveness at Mach 0.98 at 35,000 feet, less
than 2 "GY 1 are available; therefore, until the
airplane is slowed down, the elevator will have
little effect for recovery. At speeds of Mach
0.90 and above, the speed brakes will open
only 30 degrees or less, and because of adverse
compressibility effects, little drag will result
from their use. In a vertical or near vertical
dive at high Mach numbers any delay in start¬
ing recovery, combined with the greatly
reduced elevator and speed brake effectiveness,
may result in such loss of altitude that re¬
covery may be impossible. Therefore, use ex¬
treme caution in performing high Mach dives
at angles greater than 60 degrees, and make cer¬
tain that recovery from any high Mach dive
is initiated no lower than 35,000 feet. The
flight path for the 90 degree dive show r n in
figure 6-6 illustrates the excessive loss of alti¬
tude during vertical dive recovery.
Enter the dive with a wingover* Maintain positive
'‘Gy throughout the dive to prevent fiameouc. Since
in a steep dive a high percentage of the airplane's mo¬
mentum is caused by weight as compared to engine
thrust, the speed of descent can be varied only within
relatively narrow T limits by throttle changes. Observe the
effect of buffet as the airplane accelerates to high Mach
numbers and again as it decelerates during pullout. The
airplane has normal dive attitude and responds to a nor¬
mal recovery technique. Begin normal recovery pro¬
cedure at approximately 35,000 feet. See figure 6-6 for
correct procedure.
WARNING
Do not use excessive elevator trim in recover¬
ing from a dive. When airplane slows down
during pullout, elevators become more effec¬
tive, and applied trim may result in pulling
"GY* in excess of the load factor limit.
At approximately Mach 0.75, stick pressure is light
and elevators are most sensitive. Exercise caution in
this airspeed range so that design load factor is not
exceeded. Because of elevator power limits you may be
able to pull only approximately 1.3 "GY* at the begin¬
ning of recovery and about 2.5 "GY* maximum at the
end of the pullout. The exact available load factor is,
of course, dependent on Mach number and altitude.
WARNING
Since the airplane can lose altitude rapidly,
avoid steep low-level dives.
Hot©
The windshield and canopy defrost and defog
system should be operated at the highest tem¬
perature possible (consistent with aircrew
comfort) during high altitude flights. This
high temperature will keep the transparent
surfaces preheated and will preclude the for¬
mation of frost or fog during descent.
Changed 13 February 1959
6-9
Section VI
TO. 1F-S9H-1
The solid lines {elevator control power /tmifs) on
the right oj tke chart show the maximum airspeeds
at which the “GV 1 shown on the chart can
be pulled ♦ Greater speeds will result in decreased
elevator effectiveness.
Tke dotted Hues (stall limits) on the left of the
chart show the airspeed at which the airplane will
enter an accelerated stall while pulling the
“GV* sfeoim on the chart.
ALTITUDE AT START OF PULLOUT (25,000 FEET>
MOVE TO RIGHT TO AIRSPEED
AT START OF PULLOUT (300 KNOTS),
MOVE DOWN CHART TO
DIVE ANGLE CURVE (60 DEGREES),
MOVE TO LEFT AND READ FROM
THIS SCALE THE ALTITUDE LOST
DURING DIVE RECOVERY.
If airplane configuration or pou>er settings
are suck as to cause deceleration during
NO TE : dive recovery , the altitude lost will be less
than that shoum on tke charts.
6-10
Figure 6-5 (Sheet l of 6).
ALTITUDE (FEET) PRESSURE ALTITUDE (FEET)
LOST DURING RECOVERY AT START OF RECOVERY
T.O. 1F-89H-1
Section VI
AL T/TUDI LOST
DURING Dm RtCOVERY
STALL LIMITS FOR 31.677 LB GROSS WEIGHT
STALL LIMITS FOR 47,355 LB GROSS WEIGHT
ELEVATOR CONTROL POWER LIMITS FOR 31*677 LB GROSS WEIGHT
ELEVATOR CONTROL POWER LIMITS FOR 47.355 LB GROSS WEIGHT
10.000
50.000
150
300
20,000
SEA
LEVEL
2000
4000
6000!
8000
10.000
12.000
14,000
16.000
20,000
data as OF: 14 August
I I 1
_ DATA BASIS: Flight test
400
Figure 6-5 (Sheet 2 erf 6).
6-11
DIVE ANGLE (DEGREES) AIRSPEED (KNOTS IAS)
Section VI
T.O. 1F-89H-I
.v. *-W,, V»V: . • ■ ■
AL T/TUDE LOST
DURING DIVG RECOVERS
!i
V
§
*
S„
Go
$4 ^
a.
STALL LIMITS FOR 3L677 LB GROSS WEIGHT
STALL LIMITS FOR 47,355 LB GROSS WEIGHT
ELEVATOR CONTROL POWER LIMITS FOR 31,677 LB GROSS WEIGHT
ELEVATOR CONTROL POWER LIMITS FOR 47 355 LB GROSS WEIGHT
DATA AS OF: 14 August 1957
DATA BASIS' Flight test
H97 (31A
For example of chart use , see Figure 6-6,
i Sheets 1 and 2 of 6.
6-12
Figure 6-5 (Sheet 3 of 6)
DIVE ANGLE (DEGREES) AIRSPEED (KNOTS IAS)
ALTITUDE (FEET) PRESSURE ALTITUDE (FEET)
LOST DURING RECOVERY AT START OF RECOVERY
T,0* 1F-89H-1
Section VI
ALTITUDE LOST
DURING DIVE RECOVERY
STALL LIMITS FOR 31,677 LB GROSS WEIGHT
STALL LIMITS FOR 39.477 LB GROSS WEIGHT
ELEVATOR CONTROL POWER LIMITS FOR 31,677 LB GROSS WEIGHT
ELEVATOR CONTROL POWER LIMITS FOR 39,477 LB GROSS WEIGHT
ISO 200
SPEED RESTRICTION -
470 KNOTS IAS OR MACH 0.90
WHICHEVER IS LESS
4000
8000
DATA AS OF: J4 August 1957
data BASIS: Flight Test
For example of chart use , see Figure 6-6 f
Sheets 1 and 2 of 6.
H-77WA
Figure 6-5 (Sheet 4 of 6).
6-13
DIVE ANGLE (DEGREES) AIRSPEED (KNOTS IAS)
ALTITUDE (FEET) PRESSURE ALTITUDE (FEET)
LOST DURING RECOVERY AT START OF RECOVERY
Section VI
T.O. 1F-89H-1
ALTITUDE LOST
* DURING DIVE RECOVERY
STALL LIMITS FOR 31.677 LB GROSS WEIGHT
STALL LIMITS FOR 39.477 LB GROSS WEIGHT
ELEVATOR CONTROL POWER LIMITS FOR 31,677 LB GROSS WEIGHT
ELEVATOR CONTROL POWER LIMITS FOR 39,477 LB GROSS WEIGHT
150 200
SPEED RESTRICTION-
470 KNOTS JAS OR MACH 0.90
WHICHEVER IS LESS
data AS OF: 14 August 1957
DATA BASIS: Flight test
For example of chart use , see
Sheets 1 and 2 of 6 ♦
HJJ715U
6-14
Figure 6-5 (Sheet 5 of 6).
DIVE ANGLE (DEGREES) AIRSPEED (KNOTS IAS)
altitude lost
WRING Dm RECOVERS
AT CONSTANT
ACCELERATION
T.O. 1F-89H-1
Section VI
8000
DATA AS OF; 14 AuuUHt I9i>*
DATA BASIS: Flight test
For example of chart use* see Figure 6-6,
Sheets I and 2 of 6.
Figure 6-5 (Sheet 6 of 6).
10.000
40,000
30,000
20.000
10.000
SEA
LEVEL
2000
4000
6000
550
300
350
400
450
500
30
40
50
60
70
SO
90
mm STALL LIMITS FOR 31.677 LB GROSS WEIGHT
STALL LIMITS FOR 39.477 LB GROSS WEIGHT
— ELEVATOR CONTROL POWER LIMITS FOR 31.677 LB GROSS WEIGHT
ELEVATOR CONTROL POWER LIMITS FOR 39,477 LB GROSS WEIGHT
150 200
50.000
REGION RESTRICTED TO
5.00 "G" MAXIMUM——,
acceleration
SPEED RESTRICTION
470 KNOTS IAS OR MACH 0,90
WHICHEVER IS LESS
Section VI
TO. 1F-89H-!
FLIGHT WITH ASYMMETRICAL LOADING.
Flights with asymmetrical loading should be avoided if
possible* The most probable cause of asymmetrical
loading would be uneven foe! consumption between
the left and right fuel systems. If, through malfunction
or mismanagement of the fuel system, an asymmetrical
load condition develops, first attempt to correct the
condition by balancing fuel load (see Section VII) or
dumping tip tank fuel. If this cannot be done,
land as soon as practicable to preclude the possi¬
bility of the condition becoming worse. When flying
with one full and one empty tip tank, lateral control
cannot be maintained down to stall speed using trim
alone, but requires additional aileron stick force.
With trim alone, control can be maintained with full
flaps down to about 150 knots IAS. Flying near stall
speed is not recommended because nearly full aileron
deflection is necessary to maintain level flight. Land¬
ing may be made using about one-half aileron and an
airspeed above 140 knots IAS until just before touch¬
down to provide adequate lateral control.
Nofe
With clean configuration in level flight, the
airplane may start to snake through the air
at about 280 knots IAS if the sideslip stability
augmenter is not operating properly,
FLIGHT WITH EXTERNAL LOADS.
Flight characteristics (such as buffet, stall, stability,
and control) are essentially the same with or without
pylon tanks except for the restrictions covered in
Section V* Pylon tanks should be dropped before en¬
tering combat.
Hofe
H 1000)6
DOUBLE CHECK OPERATION OF ALL
CONTROL SURFACES AND HYDRAULIC SYSTEMS.
IT IS MANDATORY THAT BOTH SYSTEMS
BE OPERATING AT NORMAL PRESSURE FOR
SATISFACTORY CONTROL DURING A
HIGH MACH DIVE AND RECOVERY.
OPEN SPEED BRAKES 5
WING DROP.
TO PREVENT
ENTER 60 DIVE IN A DIVING TURN.
MAINTAINING POSITIVE “G'S" TO PREVENT
FLAMEQUT.
ENTER 90 DIVE WITH A HALF ROLL AND
MAINTAIN MAXIMUM AVAILABLE 'G'S‘*
THROUGHOUT DIVE AND RECOVERY.
ESTABLISH ANGLE OF DIVE AS SOON
AS POSSIBLE.
External stores other than pylon tanks will
not be carried.
6-16
T.O. 1F-89H-1
Section Vi
am # asm
MACH 0.80
(APPROX)
NOTE
Due to high inertia forces, initial
response to stick back pressure
is not immediately apparent.
MACH 0.90, 2.5 “GV’ (APPROX)
H-iooraw
Figure &-6,
6-17
T.O. 1F-S9H-1
Section VI 1
TABLE OF CONTENTS
Afterburner Operation .. 7-3
Fuel System Operation . .. ..* * * * - 7-3
Brake System Operation ,.*.- - 7-3
Hydraulic System Operation . . ..7-6
Canopy Jettison System.- ..- - - 7-6
ENGINE.
BURST ACCELERATION.
If conditions warrant, the engines can be burst accel¬
erated by moving the throttles rapidly to OPEN. The
engine fuel control will meter the fuel required by
the engine, and normally will not pass sufficient fuel
for excessive exhaust gas temperatures or for rpm above
100 %.
Note
During a burst acceleration from 80% rpm to
maximum power, a compressor stall may re¬
sult. This will be noted by audible pulsation,
lag in rpm, and increase in tailpipe tempera¬
ture above limits.
COMPRESSOR STALL.
Compressor stall may occur at times during engine ac¬
celeration and may be recognized by a loud rumble and
vibration in the engine and rapid rise in exhaust gas
temperature, accompanied by rpm stagnation or drop.
Compressor stall is caused by a back pressure at the
compressor outlet, which in turn is usually caused by
an exceedingly rich fuel mixture. Understallconditions,
considerably greater than normal resistance to com¬
pressor rotation is encountered, resulting in the rumble
or surge previously described. Compressor stall is most
likely to be encountered under high ambient tempera¬
ture conditions during accelerations from below 80%
rpm to higher rpm, as compressor stall is a phenomena
of acceleration only and will not occur at stabilized
power settings. Since compressor stall is most likely to
occur at approximately 80% rpm, it is recommended that
engine rpm be maintained at 85% rpm or above on final
approach until committed to landing. In addition, it is
suggested that accelerations through the 80% rpm range
be made with rapid advancement of the throttle to full
open position, in order to obtain open eyelid condi¬
tions. If compressor stall is experienced, the throttle
should be retarded to below the 80%; rpm position and
exhaust gas temperature should be allowed to drop to
normal before advancing the throttle. If engine tempera¬
ture exceeds the permissible limitation, notation of this
fact should be made in DD Form 781 after landing so
that an engine overheat inspection will be made,
EXHAUST GAS TEMPERATURE VARIATION.
Because of the wide range of ambient air temperatures
encountered at various bases where the aircraft is
operated, familiarity with the corresponding varia¬
tion in exhaust gas temperature is essential to avoid
7-1
Section VI I!
T.O. JF-89H-1
damage to the engine and assure flight safety. Ab¬
normally low exhaust gas temperature for the exist¬
ing ambient air temperature will result In a loss of
thrust This could be serious on takeoff under critical
field length conditions. In cold weather, exhaust gas
temperatures at 100% rpm are considerably lower
than in hot weather. Ir is important to check the ex¬
haust gas temperature against the rpm prior to take¬
off. If the engines are operating at military power,
the exhaust gas temperatures may decrease approxi¬
mately 65°C as the altitude increases. Using maximum
power, the exhaust gas temperatures drop a maximum
of approximately 60°C between takeoff and absolute
ceiling. There is no direct control for regulating the
exhaust gas temperature; however, temperature can
be indirectly controlled by throttle settings. Starting
the afterburner causes a slight increase in exhaust gas
temperature and a drop in engine rpm. This condition
is temporary and both temperature and rpm soon
stabilize. Refer to figure 5-3, Section V, for the run¬
way temperatures and corresponding exhaust gas tem¬
peratures to be expected at 100% rpm.
OVERTEMPERATURE VERSUS ENGINE LIFE*
The operational life of a jet engine is directly affected
by the number of hot starts and high temperature and
high rpm operations. At maximum and near maximum
performance, hot section parts are exposed to tempera¬
tures requiring their functioning at near structural
limits. The turbine wheel, in particular, is subject to
early failure when subjected to serious over tempera¬
tures or repeated slight overtemperatures because it op¬
erates with a rim temperature close to the peak of toler¬
ance for the metal from which it is manufactured. The
J35 turbine wheel has operated satisfactorily for as long
as 2000 hours at normal expected steady exhaust gas
temperature. However, an increase of as little as 15°C
under the same conditions will appreciably reduce the
turbine wheel life. Transient temperatures that exceed
maximum allowable for as little as two seconds can
render the turbine wheel unserviceable. Obviously, any
overtemperatures, even momentary, beyond the limi¬
tations stipulated in Section V are serious and should
be recorded accurately. When the engine is properly
adjusted, the exhaust gas temperature indicating sys¬
tem properly calibrated, and the engine controls prop¬
erly handled, all operating temperatures including
transients will fall within the serviceability limits es¬
tablished for the engine. The careful monitoring of
exhaust gas temperature by the pilot, and the recording
of all overtemperature operation is imperative. Particu¬
larly during starting the pilot should, with a clear
understanding of the fuel flow characteristics and
their relation to exhaust temperature, be alert for an
incipient overtemperature condition and recognize
it in time to take rapid corrective action.
ENGINE OVERSPEEDING AT ALTITUDE.
The engine will operate at sea level, with or without
afterburning, within the limits preset on the engine
fuel control. However, when operating at altitude, the
fuel requirements without afterburning are somewhat
reduced and there is a possibility that the engine may
overspeed. Under most conditions the governor will
prevent the engine from exceeding 100% rpm, but be¬
cause of the inherent acceleration lag of the engine
fuel control governors, a slight engine overspeeding in
excess of 100% rpm may occur. In the event of over¬
speeding, retard the throttle to a setting that will pre¬
vent exceeding a stabilized rpm of 100%).
EYELID OPERATION*
The eyelids are provided to increase the diameter of
the tailpipe nozzle during afterburning. This is to per¬
mit an increase in thrust without operating at prohibi¬
tively high exhaust gas temperatures. In addition to
opening in conjunction with afterburning, the eyelids
will stay open during starting to prevent high tempera¬
tures, and during rapid acceleration to decrease accelera¬
tion time. An open-throttle switch and an idle switch,
both operating on 28-volt dc, are in the No. 4 inlet duct
island and are mechanically actuated by the throttle
shaft. The idle switch is actuated when the throttle
is at IDLE or below and causes the eyelids to stay open
in this speed range. The open-throttle switch is actu¬
ated when the throttle is full open and causes the eye¬
lids to open during burst accelerations, or when the
throttle is opened faster than engine rpm rises; how¬
ever, an engine speed-sensing switch will open, inter¬
rupting the open-throttle switch circuit when the
engine rpm reaches 87.5% and causing the eyelids to
close. If afterburning is selected during burst accelera¬
tion (by lifting the fingerlifts), the eyelids will stay
open during the engine speed range from idle to 100%
rpm (or from that rpm at which the burst acceleration
is started). A pressure switch is in series with the idle
switch and will open the idle switch circuit at 10,000-
foot altitude and cause the eyelids to stay closed during
high altitude idle. When the throttles are opened
slowly, the eyelids will remain closed from idle to 100%
rpm since the speed switch will be actuated in advance
of the open-throttle switch to maintain closed eyelids
during slow acceleration. Failure of the engine speed¬
sensing switch or loss of power from the primary a-c
single-phase bus, will cause the eyelids to open during
nonafterburning operation if the open-throttle switch
is closed (throttle at 100% rpm position) and the air¬
plane is below 10,000-foot pressure altitude (altitude
switch closed): This will result In an extreme loss of
thrust and low exhaust gas temperature. However, the
eyelids can be closed by moving the afterburner control
circuit breaker to the OFF position or, if trouble is
caused by failure of a single-phase inverter, by moving
the single-phase inverter switch to the EMER position.
7-2
T.O. TF-89H-1
Section VIS
The eyelids are operated by two pneumatic cylin¬
ders powered by air from the 11th stage engine compres¬
sor* The compressor air is directed to either side of the
pneumatic cylinders by a solenoid valve which is con¬
trolled by a pressure-differential switch which senses
pressure changes in the engine tailcone. If the eyelids
fail to open when afterburning is selected, engine rpm
will drop and exhaust gas temperature will rise. If this
occurs, afterburning must be discontinued immediately
to prevent excessive exhaust gas temperature, A failure
of both single-phase inverters during afterburning will
have no effect on engine and afterburner performance
until the afterburners are shut down* If the airplane is
below 10,000 feet and the throttles at 100% rpm, the
eyelids will have to be closed by moving the after¬
burner control circuit breaker to the OFF position.
Afterburning will not be available again until the
single-phase power failure is corrected* Failure of the
eyelids to dose following afterburner operation will
result in very low exhaust gas temperature and extreme
loss of thrust.
AFTERBURNER OPERATION.
STARTING AFTERBURNERS AT HIGH ALTITUDE.
If difficulty is encountered when initiating after¬
burning at altitudes above 45,000 feet using the nor¬
mal procedure, use the following procedure to decrease
the time required to reach full afterburner operation.
L Retard throttle to 95% rpm*
2. Lift throttle fingerlift, and simultaneously jab
the throttle forward. Large jabs of more than 3% rpm
are not recommended as they may result in overtempera¬
ture conditions*
FUEL SYSTEM OPERATION.
See figure 7-1 for fuel flow during normal sequencing
and figure 7-2 for fuel flow during manual selection of
wing tanks and cross feed operation,
CROSSFEED OPERATION.
A 28-vok d-c crossfeed switch (figure 7-1), located on
the fuel control panel, has OPEN and CLOSED posi¬
tions. When the crossfeed switch is at OPEN, the main
fuel lines of both systems are interconnected; both fuel
systems may be used to operate one engine or both en¬
gines may be operated from either fuel system* Unbal¬
anced lateral fuel loading (wing heaviness) may be cor¬
rected by feeding both engines from the system having
more fuel. To balance fuel load, place the crossfeed
switch at OPEN and the fuel selector switch for the
system with less fuel at PUMPS OFF. When fuel load is
balanced, as indicated by lateral trim and/or fuel quan¬
tity gages, return the selector switch to ALL TANKS,
and the crossfeed switch to CLOSED. With one engine
inoperative and the crossfeed switch at OPEN, fuel will
be supplied to the operative engine from both fuel
systems in either ALL TANKS or WING TANKS
selection.
The throttle for the inoperative engine should
be closed. If the throttle is left open, the
throttle controlled fuel shutoff valve will be
open allowing fuel to be metered through the
engine.
BRAKE SYSTEM OPERATION.
Wheel brakes should be properly used and treated
with respect to reduce maintenance difficulties and ac¬
cidents due to wheel brake failure* Brakes should not
be dragged when taxiing and should be used as little
as possible for turning the airplane on the ground* Ex¬
treme care should be used to prevent locking a wheel
and skidding the tires when applying brakes immedi¬
ately after landing when there is considerable lift on
the wings. Proper brake action does not occur until the
tires are carrying heavy loads. Heavy brake pressure
can result in a locked wheel far more easily if brakes
are applied immediately after touchdown than if the
same pressure is applied after the full weight of the
airplane is on the wheels. Brakes can stop a wheel from
turning, but stopping the airplane is dependent on
the friction of the tires on the runway. Skidding re¬
sulting from improper braking tears off shreds of rub¬
ber that act as rollers between tire and runway; the
heat generated by skidding melts tire rubber and the
resultant molten rubber acts as a lubricant between tire
and runway. The full landing roll should be utilized to
minimize the use of wheel brakes and to take advantage
of aerodynamic braking* Using either normal or emer¬
gency braking systems, short landing rolls (executed
only when necessary) are accomplished by a single,
smooth application of brakes wirh constantly increas¬
ing pedal pressure. To allow sufficient time for cooling
between brake applications, a 15-minute interval is
required between full stop landings where the landing
gear remains extended in the slipstream and 30 minutes
between full stop landings where gear has been re¬
tracted. If the brakes are used for steering or cross-
wind taxiing, or if a series of landings is performed,
additional time for cooling is required. When the
brakes are in a heated condition resulting from ex¬
cessive use in an emergency stop, the airplane should
not be taxied into a crowded area and the parking
brake should not be set. Peak temperatures occur from
5 to 15 minutes after a maximum braking operation
and proper brake-cooling procedure should be fol¬
lowed to prevent brake fire and possible wheel as¬
sembly explosion. On airplanes modified in accord¬
ance with T.O, 1E-89H-522, an antiskid braking de¬
vice is incorporated in the brake system. This device
is designed to allow maximum braking efficiency
during normal and adverse weather conditions with¬
out skidding the main wheels.
7-3
Section VII
T.O. 1F-89H-T
Emptied fuel spare
TIP AND PYLON TANK FUEL FLOW
normal roti smiNcme
WING TANK FUEL FLOW
l * 1 'OOOCC
.Tl
Uis« \yw i
k! sf#-*.
S^s.s.-s'S 5aisr
> ">oooo<'r
Figure 7-1 *
7-4
T.O. 1F-B9H-1
Section VII
Figure 7-2.
7-5
Section VII
TO, 1F-B9H-1
HYDRAULIC SYSTEM OPERATION,
Hydraulically powered systems whose normal opera¬
tion is standard to most aircraft will not be discussed
in this section,
WING FLAP OPERATION.
The wing flap lever can be pre-positioned at UP,
TAKEOFF, or DOWN; and the flaps will move to
the selected position* For intermediate positions, the
lever most be held at the desired position until the in¬
dicator shows the flaps to be in that position. The
lever can then be released and the flaps will remain in
position until the lever is moved,
SPEED BRAKE OPERATION.
The speed brake lever opens the speed brakes propor¬
tionately to the lever movement. Pre-positioning the
lever at any point toward the OPEN limit of travel will
stop the speed brakes in the corresponding posi¬
tion, At indicated airspeeds up to approximately 260
knots, the speed brake surfaces can be opened to any
position (from 0 degrees to 120 degrees included angle).
At indicated airspeeds above 260 knots, the angle to
which the speed brakes open will be decreased propor¬
tionately to the increase in airspeed. If the airspeed is
great enough, the airflow creates a back pressure in the
system and the speed brakes will "blow back” to the
point where the back pressureon the actuating cylinders
is equal to that of a relief valve in the speed brake hy¬
draulic line* As the airspeed decreases, the speed brakes
open to the original position if there has been no change
in the position of the speed brake lever* The speed
brake cannot be pre-positioned toward the CLOSED
position. The speed brake must be pushed forward
manually as the speed brakes close.
CANOPY JETTISON SYSTEM.
To properly jettison the canopy, a minimum pressure
of 1400 psi is required in the canopy jettison system*
The decrease in temperature which accompanies high
altitude flight may cause the cylinder pressure to drop
below 1400 psi* A pressure of 1800 psi when the am¬
bient temperature is 1O0°F (38°C) will assure a mini¬
mum pressure of 1400 pst if the temperature decreas¬
es to — 50°F ( — 46°C)* To determine the required
pressure at other ambient temperatures, subtract 40
psi from 1800 psi for each 15°F (8*4°C) decrease
below 10O°F (3S°C), For example, the required pres¬
sure for an ambient temperature of 70°F (2I°C) would
be 1720 psi*
7-6
T.O. 1F-89H-1
Section Vii!
TABLE OF CONTENTS
Pilot's Duties . . . ..*... 8-1
Radar Observer's Duties ..8-1
Abbreviated Checklist . 8-5
CREW DUTIES
PILOT’S DUTIES.
The duties of the pilot have been covered thoroughly
in other sections of this handbook and will not be re¬
peated here.
RADAR OBSERVER’S DUTIES.
The radar observer's primary duty is to operate the
radar equipment; therefore, he must be on every mis¬
sion in which the radar equipment will be used. In
addition to operating the radar equipment, he reads all
checklists to the pilot and performs other important
duties which are covered in the following paragraphs.
Note
For reasons of security classification, informa¬
tion concerning the E-9 fire control system
and armament is not included in this manual.
For information covering this equipment,
consult T.O. 1F-89HTA,
EXTERIOR INSPECTION,
At the discretion of the pilot, the radar observer will
assist in making the exterior inspection (figure 2-1).
;■ CAUTION J;
& » #+** # ##»+** +#*** » # *
On some airplanes, two lockbolt position
indicators on each engine nacelle door are
provided to permit visual reference of their
position when doors are being locked. When
the small inspection door cover plates are
removed, a movable lockbolt position indi¬
cator and a stationary reference indicator will
be visible. These indicators must be aligned
within 1/32 inch when the lockbolt is in
locked position.
BEFORE ENTERING COCKPIT.
1. Ejection seat—Check:
Armrests and trigger stowed; safety belt release
initiator ground safety pin—Removed; safety
pins installed; catapult file mark aligned.
Note
If the safety belt initiator ground safety pin is
installed, consult maintenance personnel re¬
garding the status of the ejection system be¬
fore occupying the ejection seat.
2. Flashlight—Check operation.
3. Circuit breakers—In.
ON ENTERING COCKPIT,
Note
A radar observer’s checklist is located on the
radar observer's instrument panel.
Changed 13 February 1959
8-1
Section VIII
T.O 1F-89H-1
INTERIOR CHECK,
Rear Cockpit.
If the C-2A life raft is being carried, the A-5
seat cushion should not be left on the seat. If
both are used, and it becomes necessary to
eject or crash land, severe spina! injury may
result due to the excessive compressibility of
the combination of life raft and cushion. If
additional height in the seat is needed, a solid
filler block may be used in conjunction with
the life raft.
Note
When the seat cushion is not used, the Type
MD-i contoured seat style survival kit con¬
tainer, stock number 2010-126602, with the
MA-1 contoured cushion, stock number 2010-
159100, should be used. The forward edge of
the packed kit should not be thicker than 7
inches (consult T. O. 14S1-3-51, "Base As¬
sembly, Use and Maintenance of Sustenance
Kits” and T.O. 14S3-2-31, "One Man Life
Raft, Type PK-2, Used with Survival Kit
Container, Type MD-1”), The CA-2 one-man
life raft may be used if the MD-1 containers
are not available.
L Safety belt and shoulder harness—Fasten; inertia
reel operation—Check; static cord lanyard—Con¬
nected; automatic-opening parachute lanyard—Con¬
nected,
• If the safety belt is opened manually, the
parachute ripcord must be pulled manually.
• Improperly attaching the shoulder harness
and safety belt tie-down straps to the auto¬
matic belt will prevent separation from the
ejection seat after ejection. To make the at¬
tachment correctly, first place the right and
left shoulder harness loops over the manual
release end of the swivel link; second, place
the automatic parachute lanyard anchor over
the manual release end of the swivel link;
then, fasten the safety belt by locking the
manual release lever.
• The M-4 or M-12 safety belt initiator ground
safety pin with the warning streamer must
be removed prior to flight. If the pin is not
removed, automatic uncoupling of the safety
belt will not occur if ejection becomes neces¬
sary. If pin is installed, maintenance person¬
nel should be consulted on the status of the
ejection system before the seat is occupied.
2, E-9 fire control test panel—Check (see T.O. IF-
89IMA).
3- Alternator breaker control switch momentarily at
TRIP; external power switch—CLOSE (after external
power is connected).
4. I n te rphone a mpl if ie r switch—-ON.
5. Interior light switches—As necessary.
6. Canopy defog knob—IN.
7. Altimeter and clock—Set and cross-checked with
pilot.
8. Canopy jettison pressure gage—Check pressure.
9- Interphone selector switch—COMM INTER; in¬
terphone toggle switch—INTER.
10. Communications equipment—Check operation.
11. Emergency signal system—Check (with pilot).
12. Oxygen equipment—Check operation.
Pressure gage 400 psi; oxygen regulator diluter
lever NORMAL OXYGEN; oxygen regulator
supply lever ON. (Refer to Oxygen System Pre-
flight Check, Section IV, for detailed informa¬
tion.)
13. Hydraulic handpump system—Check.
Engine hoist and brake selector valve handles
positioned with aft handle (B) to NEUTRAL
and forward handle (A) to SYSTEM; hand-
pump handle stowed.
Note
For additional instructions regarding the ra¬
dar observer’s equipment, refer to T.O. IF-
89H-1A.
GROUND TESTS,
1. II 5-volt alternator system—Check.
With left engine rpm above 60%, move alter¬
nator exciter switch and alternator circuit
breaker switches to CLOSE momentarily. Check
alternator voltmeter for 11> ±r 1,5 volts.
2. Inverter buses—Check voltage.
Check both single*phase inverter buses and
three-phase bus for proper voltage; recheck
voltage of three-phase bus when pilot selects
spare instrument inverter,
BEFORE TAKEOFF,
1. Ejection seat ground safety pins—Remove.
2. Safety belt—'Lighten; shoulder harness—Adjust
to fit snugly; inertia reel lock lever—LOCKED.
3. Anti "G” suit valve button— Press to check opera¬
tion.
AFTER TAKEOFF—CLIMB.
1. Static cord lanyard—Disconnected above mini¬
mum safe ejection altitude.
8-2
Changed 13 February 1959
T.O, 1F-89H-1
Section VIII
DURING FLIGHT*
1. Adjust radar controls for set operation (see T,G,
1F-89H-1A).
BEFORE LANDING*
1. Safety belt and shoulder harness-—Check for tight¬
ness; static cord lanyard—Connected above minimum
safe ejection altitude.
2. Viewing scope—Place in stowed position,
3. Radar console assembly-—Move to forward posi¬
tion,
4 . Inertia reel—LOCKED,
BEFORE LEAVING AIRPLANE*
L All switches—OFF,
2, Ejection seat ground safety pins—IN.
Note
The following checklist is an abbreviated ver¬
sion of the procedures presented in the simpli¬
fied checklists of Section VIII. This abbrev¬
iated checklist is arranged so you may remove
it from your flight manual and insert it into
a flip pad for convenient use. It is arranged
so that each action is in sequence with the
amplified procedure given in Section VIII,
8-3
T.O. 1F-89H-1
Section VIII
ABBREVIATED CHECKLIST
CUT ON DOTTED LINE
NORMAL PROCEDURES
F-89H ABBREVIATED CHECKLIST
(Radar Observer)
Note
The following checklist is an abbreviated version of the radar
observer’s duties and is accomplished by the radar observer.
EXTERIOR INSPECTION
At the discretion of the pilot, the radar observer will assist in making the
exterior inspection (figure 2-1).
BEFORE ENTERING COCKPIT
1. Ejection seat—Check.
2. Flashlight—Check operation.
3. Circuit breakers—IN.
INTERIOR CHECK
REAR COCKPIT
1. Safety belt and shoulder harness—Fasten; inertia reel operation—
Check; static cord lanyard—Connected; automatic-opening para¬
chute lanyard—Connected,
2. E-9 fire control test panel—Check (see T.O. 1F-89H-IA).
3. Alternator breaker control switch momentarily at TRIP; external
power switch—CLOSE (after external power is connected),
4. Interphone amplifier switch—ON.
5. Interior light switches—As necessary.
6. Canopy defog knob—IN.
7. Altimeter and clock—Set and cross-checked with pilot.
8. Canopy jettison pressure gage—-Check pressure,
9- Interphone selector switch—COMM INTER; interphone toggle
switch—INTER.
10, Com m u nicad o ns equ i pme n t—Check ope ra tion,
T.O, IF-89H-1 1
31 OCTOBER 1958
CONTINUED ON NEXT PAGE
8-5
Section VIII
T.O. 1F-89H-1
ABBREVIATED CHECKLIST
CUT ON DOTTED LINE
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-uopujado spaq;)—suoiudrnba ua^XxQ ’Zl
*(TO[id qjiAi) spaq;)— ujzisAs puSis iouagjauij y\
T.O. 1F-89H-1
Section IX
SECTION IX
AU-WEATHER OPERATION
The procedures in this section pertain only to all-weather operation and are in addition to the nor¬
mal procedures in Sections II and IV. Normal procedures are repeated here only where necessary.
TABLE OF CONTENTS
Page
Instrument Flight Procedures ..*..... 9-1
Ice and Rain ..... 9-13
Turbulence and Thunderstorms . , ...... 9-T5
Night Flying .. . , . .... ....... 9-16
Cold Weather Procedures ....* , • ..*.* , . ..9-T6
Hot Weather Procedures.. . ..9-20
Desert Procedures . . . . ......... . . . 9-21
Except for some repetition necessary for emphasis, clarity, or continuity of thought, this section
contains only those procedures that differ from or are in addition to the normal operating instruc¬
tions covered in Sections II and IV relative to instrument flight.
INSTRUMENT FLIGHT PROCEDURES
INTRODUCTION.
Flying the airplane in instrument weather conditions
requires instrument proficiency and thorough pre-
flight planning. In planning IFR flighrs, remember
that fuel requirements for completion of instrument
letdown approach procedures and possible diversion to
alternate fields must be added to that normally re¬
quired for VFR flights. Therefore, maximum range
or endurance of the airplane, if required to land
in IFR weather conditions, is reduced accordingly.
The airplane has good stability characteristics and
flight handling qualities for weather flying. For ease
of handling, banks should be limited to 30 degrees
unless maximum rate turns are ordered by GCI during
interceptions. The flight computer installation great¬
ly simplifies precision instrument flying. Pilots should
avoid any tendency, however, to concentrate exclu¬
sively on the flight computer indicator or to be
hypnotized by it . Concentration on the indicator alone,
particularly during rollout from turns, may cause a
temporary sense of vertigo . When using the flight
computer f monitor the action of the airplane with the
basic standard flight instruments at all times to be
sure that the airplane follows the flight path set up on
the flight computer controls .
Section IX
TO, 1F-B9H-T
NOTE:
Cross-check with all basic
flight instruments th rangfum t
takeoff to determine proper
flight attitude-
INSTRUMENT TANEOfF
WITH FUCNT COMPUTER (Typical)
TAKEOFF
A* Taxi into position and make visual lineup
on center of run tray*
Ik Set course dial on flight computer indicator
to coincide with runway heading.
LIFTOFF
lift the airplane off the runway in normal
manner amt zero the horizon tal Imr. The (wo
dots IK-a p setting will automatically provide
a sale amt efficient takeoff and initial
rlimh to a <afe terrain altitude.
BEFORE TAKEOFF
A. Flight computer selector switch—FLIGHT INST.
£L Set horizontal bar of flight computer
indicator nl two dots IK-up signal*
3
GROUND ROLL
Maintain heading with nose wheel ^Leering
until the rudder becomes effective (approx,
(»<) knots IAS). Mold vertical bar an center.
H-1D3C
1
Figure 9-T,
9-2
TO, 1F-89H-1
Section IX
1 INITIAL CLIMB
Ai a safe altitude above terrain,
accelerate to be&r climbing ainpe^d.
NOTE: Cross-check t nth basic
flight instruments durine climb
i 'I | \i i) i)
and after leveling off. ULiiJin
fCsUihUsh flcKiml illicit- of dimli ami
adjust horizontal bar to sera with
the pi tell trim knob.
cum WITH fUGHT COMPUUR
(Typical)
LEVELING OFF 4
When the desired altitude Is readied,
turn altitude control switch to ON' and
Kero the horizontal but*. Return the
jutell trim Limb to its normal horizontal
position (knob pojnlio(tto index murk)
3
Keep horizontal bur wto« 1 at best climbing
airspeed by reducing the pitch trim os
neeeuaary during climb to allRude.
H-104C
Figure 9 - 2 .
Section IX
T.O. 1F-89H-1
INSTRUMENT TAKEOFF.
Instrument takeoffs without afterburning are not rec¬
ommended. Afterburning is recommended to shorten
the takeoff roll in conditions of low visibility and
when takeoff in cross wind is made. After completing
the prescribed Taxi and Before Takeoff checks and
after aligning the airplane on the runway, set the
course dial on the flight computer indicator to coin¬
cide with the runway heading. As the takeoff roll is
started, maintain proper directional control with nose
wheel steering until the rudder becomes effective at
approximately 70 knots IAS. Maintain heading with
reference to the directional indicator. Concurrent use
of runway markers and visual references, as long as
they remain visible, is recommended. Continue the
instrument takeoff, lifting off the nose wheel and
becoming airborne at the normal VFR speeds. Estab¬
lish and maintain the proper attitude on the attitude
indicator until definitely airborne. As the airplane
leaves the ground the attitude indicator is primary
for both bank and pitch and remains primary until
the climb is definitely established. When the ver¬
tical velocity indicator and the altimeter show a defi¬
nite climb indication, retract the gear and flaps as
under VFR conditions. Upon reaching a safe altitude,
accelerate to a normal climb speed. If necessary, turn
the anti-icing switch to FLIGHT.
Note
Approximately 5 degrees of roil error may
appear on the attitude indicator on acceler¬
ated turn after takeoff. This error will be in
the direction of the turn and should disap¬
pear within a short time. See figure 9-1 for
typical instrument takeoffs with the flight
computer.
INSTRUMENT CLIMB*
Once the desired climb speed is reached the airspeed
indicator becomes the primary instrument for pitch
and remains as such throughout the remainder of the
climb. Refer to figure 9-2 for a typical flight computer
climb. Use the climb procedures as outlined in Sec¬
tion II.
INSTRUMENT CRUISING FLIGHT*
After leveling off and adjusting power as necessary,
trim the airplane for hands off flight. Altitude may be
maintained by holding the horizontal bar of the flight
computer centered, with the altitude control switch
ON. However, the altimeter is still the primary instru¬
ment for pitch, since only it can provide the pilot
with an indication of altitude. The attitude indicator
is the only direct reading instrument for pitch and
bank changes. Turn errors occur in both its pitch and
hank indications. Asa result a close cross-check on the
altimeter and turn needle must be accomplished in
rolling out of turns. After a short time the gyro will
precess back to a correct indication. In accomplishing
turns with the flight computer the maximum bank
angle required to center the vertical bar is set at 30
degrees regardless of airspeed and altitude. Banks of
more than 30 degrees may be made by holding the
vertical bar at one or more dots beyond center. The
maximum amount of heading change that should be
selected on the flight computer at any time is 150 de¬
grees. If when flying on a heading of 360 degrees a
right turn to 180 degrees is desired, rotate the heading
selector until 150 degrees is under the course index.
Start the turn, and when more than 30 degrees of the
turn have been accomplished, rotate the heading se¬
lector to 180 degrees and continue the turn. The flight
computer will initiate a rollout indicating 22 degrees
before the selected heading is reached. It is advan¬
tageous to roll out within reference to the vertical bar
when a more rapid change of heading is desired.
Note
If more than 150 degrees from present head¬
ing is selected under the course index, the
vertical bar will indicate a turn in the oppo¬
site direction.
See figure 9-3 for typical flight computer mm proce¬
dure.
IFR INTERCEPTIONS.
With sufficient practice, interceptions can be per¬
formed under instrument conditions without difficul¬
ty. With proper coordination between pilot and radar
observer, the pilot can perform the attack phase of the
interception under instrument conditions, using the
attitude indication and target information on his ra¬
dar scope. Use of the flight computer in conjunction
with the E-1I autopilot during the initial phase of an
intercept when under GCI control greatly simplifies
instrument flight during ground control phase of inter¬
ceptions. When given vectors by the GCI controller,
turn the flight computer heading selector to the corres¬
ponding heading, and roll immediately into the turn
to center the vertical bar on the indicator. Keep the air¬
plane trimmed while tracking the target, particularly
when decelerating after lockon. The attack phase can
be flown by the attitude reference presented on the
radar scope. Use the windshield wiper in precipitation
to increase visual sighting range after lockon.
9-4
■VOTE: When using the flight campnter,
crass check zrif/i basic flight instruments.
TOUHS WITH TUCHT COMPUTCR
FUCHT
I'KST.
JfFL ■
AP-PROACM
r APPROACH
fm
\ OFF
■tsooe
, S-WITCH
ji' m+::+ mh
flfcHl
EmT
APPPQACh
LECTOR M* SWITCH
Ah-PROACH
lector iMk Switch
When ilir iwuf ullihidp in approadiHl, level
niT with Tpffppjirr to the liasie fit till l instrument*,
Vfler airplane is levrffd off at tin- new altitude
turn I In- altitude ronlml rmilrh lit ON and kee
the horizontal har centered, Ketiirii the pilch
trim knob to it* normal horizontal position
( knoh pointin'! to index mark).
Figure 9-4
T.O. TF-89H-1
Section IX
SPEED RANGE.
Airplane flight characteristics at high and law air¬
speeds are the same for VFR and IFR flying. For best
cruise or loitering indicated airspeeds refer to appli¬
cable Appendix charts.
RADIO AND NAVIGATION EQUIPMENT.
For proper background and use of radio and naviga¬
tion equipment refer to Section IV, The operation of
radio and navigation equipment is not affected by most
weather conditions. The radio compass* however, is
susceptible to precipitation static.
DESCENT.
If icing conditions are probable, the descent should be
made with sufficient power to provide adequate hot
air for the anti-icing system. For maximum ease of
handling, a constant-speed letdown is recommended.
The optimum speed brake position depends on the
IAS and rate of descent combination desired. The
adjustable speed brakes make possible various rates of
descent at the same IAS and throttle setting.
RADIO PENETRATIONS.
Radio penetrations can be accomplished satisfactorily
with various airplane configurations. Recommended,
however, is an 85% rpm, 250-knot IAS and 4000 fpm
descent, maintained with gears and flaps retracted and
approximately one-half speed brakes. The exact proce¬
dures for jet penetrations (Pilot's Handbook — Jet, East
or West) will vary with each field due to local terrain
and radio variations.
Note
The canopy defogging system should be actu¬
ated approximately 10 minutes prior to de¬
scent from altitude.
See figure 9-4 for a typical flight computer descent
procedure.
INSTRUMENT APPROACHES.
The airplane has excellent handling characteristics
during instrument approaches. When power is at idle
or low rpm the power response to throttle movement
is very slow. Therefore, use relatively high power set¬
tings in the approach configuration, and control air¬
speed and rate of descent by using the speed brakes.
Very little pitch change is required during transition
from glide slope to touchdown, because the airplane is
approximately in a landing attitude while on the glide
slope. With flaps at takeoff, speed brakes open, and
maximum practicable braking, the required runway
length to stop, following instrument approaches, is
short compared to other jet fighters. A 65 00-foot
GCA or ILS equipped runway is considered minimum
for actual all-weather operations.
RADIO APPROACHES.
Normally, radio range and omnirange approaches
will be required only if the airplane is not VFR after
descent to the low station and no GCA or ILS is avail¬
able, Refer to the Pilot's Handbook—Jet for the local
procedures of the standard instrument approach. The
fuel required to complete an approach is largely deter¬
mined by the time the airplane flies outbound before
making the procedure turn and by the distance from
the fix to the field. The time outbound from the radio
fix, prior to initiating the procedure turn, need only
be sufficient to permit completion of the cockpit check
after the procedure turn and precision beam following
to the station at the proper altitude. For radio ap¬
proaches after a tear-drop type penetration the follow¬
ing procedures may be used.
Note
If a procedure turn is to be made after a
penetration, use 85% rpm and adjust speed
brakes as required to maintain 195 knots IAS.
Fly outbound for a minimum of 30 seconds
and a maximum of 60 seconds (or as locally
prescribed); then make procedure turn.
INBOUND.
1. Landing gear lever—DOWN.
2. Wing flap lever—TAKEOFF.
3. Throttle—Minimum of 85% rpm.
4. Speed brake lever—As required to maintain 160
knots IAS,
5. Descent to proper altitude.
Note
If the time from the radio fix to the field ex¬
ceeds 2 minutes, it is best to delay final con¬
figuration until over the station in order to
expedite the approach and conserve fuel.
LOW STATION.
Make the proper position report and descend to mini¬
mum altitude. Use the speed brakes to maintain air¬
speed in the descent. Descents during approaches are
normally made at 500 fpm and should not exceed
1000 fpm. See figure 9-6 for typical radio approach.
GROUND CONTROLLED APPROACH
(GCA).
GCA approaches may consist of a rectangular pattern,
a straight-in approach from the penetration, or modi¬
fied versions of either dependent upon local facilities
9-7
RADIO PEHITRAVQH (Typical)
1 APPROACH TO STATION
A. Canopy defo AS REQUIRED.
B. Windshield heat-AS RETIRED.
C. Pilot heat-AS REQUIRED.
O. Interior cockpit lighting—AS REQUIRED.
A, Raie-of-deseenl—Decrease I tH)0 feet
above levei-off altitude.
B. Lead level-off altitude by approximately
10% of rale of descent.
2 PENETRATION ENTRY
A. llirottlc-85% RPM.
B. Establish 4<MKt feet per mt utile
rale of descent.
C. Speed brake*—Adjust to maintain
250 KNOTS IAS.
3 PENETRATION TLRN'
Maintain descent criteria and turn
as prescribed by the appropriate
^Pilot's Handbook—Jet. ^
5 INBOUND
A. TliruUht—85% RPM.
B. Speed brakes— Ad just to maintain
2fK> KNOTS IAS.
NOTE;
Refer to appropriate "f’Mot's Handbmik — Jet*
for specific penetration instructions . ise
the basic instruments and cross-check
with the flight computer .
H407C
Figure 9 - 5 ,
9-8
TO. 1F-89H-1
Section IX
NOTE:
Refer to Pilot's Handbook for instrument
approach procedure.
The time required to perform a standard
radio range approach is approximately 10
minutes , the fuel expended, approximately
900 pounds.
RADIO APPROACH ( Typical)
1 outbound 3
A. Tbn»tlle-S5% RPM minimum,
B, Speed brake*—A* reijuired l« maintain 195 knoli* IAS.
C* Time— As lot-ally refpiired .
2 PROCEDURE TURN
4
5
COCKPIT CHECK
A. Laudiitir /rear—DOWN,
R. Wilts flap*—TAKEOFF.
C. ThroUk-85% RDM minimum *
D* Speed brakes— A* required to
in ail (tain 160 knots IAS*
INBOUND
A. Descend In proper all it tide,
B, Maintain final configuration.
LOW STATION
A* Make proper position report .
R- Descent to minimum altitude*
H 10SC
Figure 9-6 *
Section IX
T*0* TF-89H-1
and terrain features. Therefore, the fuel and time re¬
quired for a GCA will vary at different fields. The
basic procedures remain the same for all patterns* That
is, the cockpit checks and the final configuration are
accomplished prior to being turned over to the final
controller* On a cross-country flight, the GCA proce¬
dures at the destination should be checked and fuel
allowances made as part of the preflight planning*
Emergency GCA approaches can be made using less
fuel by requesting the GCA controllers to shorten the
pattern* Fuel can also be conserved by delaying the
final configuration. The procedures for a typical GCA
pattern are outlined in figure 9-7* Single-engine GCA's
can be accomplished satisfactorily using the following
procedures*
1* Downwind—T95 knots, throttle as required (ap¬
proximately 86% rpm), gear up, flaps up, speed brakes
closed,
2* Base leg—ISO knots IAS, throttle as required (ap¬
proximately 95% rpm), gear down, flaps up, and speed
brakes closed.
3* On final approach prior to glide slope entry—160
to 170 knots IAS, throttle as required (approximately
98% rpm), gear down, flaps at takeoff, and speed
brakes closed.
4* Glide slope—160 to 170 knots IAS, maintain as
high rpm as possible,
5, As end of runway is approached, do not reduce
airspeed below 160 knots IAS until landing is assured.
6* Retard throttle to idle only when positive of land¬
ing, After touchdown open speed brakes to reduce
ground roll*
ILS APPROACHES.
ILS is very similar to GCA in that it is designed to
give indications of both azimuth and elevation to the
pilot throughout the complete approach* It does differ
from a GCA since ILS gives a visual presentation of
deviations from the approach, while in GCA the pilot
is given verbal corrections throughout the approach.
The procedures for the airplane are very similar for
both GCA and ILS, and are as follows:
OUTBOUND,
L Landing gear—Up,
2. Wing flaps—Up,
3. Throttles—85% rpm minimum.
4. Speed brakes—As required to maintain 195 knots
IAS*
5* Altitude as locally required*
PROCEDURE TURN,
1* Begin procedure turn as locally prescribed*
INBOUND TO OUTER MARKER.
I* Descend to proper altitude,
2* Landing gear—Down*
3* Wing flaps—Takeoff,
4* Throttles—85% rpm minimum, to maintain l 60
knots IAS*
OUTER MARKER AND INBOUND ON APPROACH.
1, Make the appropriate position report.
2, Intercept and bracket the glide slope, maintain¬
ing airspeed with use of the speed brakes.
3- Heading corrections should not exceed 5 degrees.
Pitch corrections of 200 to 300 FPM generally will be
sufficient.
The flight computer greatly simplifies the initial turn¬
on to the localizer as well as precision beam following
on the localizer and glide slope. To use the flight com¬
puter the pilot must accomplish the steps as outlined
in figure 9-8.
ILS—AUTOPILOT-CONTROLLED
APPROACH.
Engage the autopilot and, using any standard ap¬
proach, maneuver the airplane to intercept the local¬
izer beam at approximately 45 degrees, 10 miles out,
and at 1200 to 1500 feet above the terrain. (The al¬
lowable intercept angle is 45 degrees at 8 miles, in¬
creasing proportionally to 90 degrees at 13 miles,) Use
the following general procedure to obtain consistently
good results*
1* Approach to localizer—Lower flaps and landing
gear, adjust power for 160 knots IAS, and check that
both flags on the course indicator are down. Trim the
airplane for approximately level flight at 1200 to 1500
feet above the terrain, and place the altitude switch at
ON if desired,
2. Intercepting localizer—When the airplane enters
the localizer beam, the vertical bar on the course indi¬
cator will leave its stop. As soon as this occurs, place
the localizer switch at ON* The airplane will bracket
the beam automatically,
3. Intercepting the glide slope—When the airplane
enters the glide slope, the horizontal needle of the
course indicator will approach the center of the meter*
When the needle enters the top half of the small
circle, set the approach switch at ON* The airplane
will start down the beam automatically*
4. On the glide slope—Adjust flaps and speed for
flareout and landing*
5* Breakthrough or minimum altitude—Disengage
the autopilot, complete flareout, and land manually*
9-10
T.O, 1F-89H-1
Section IX
NOTE:
The. time required to perform a standard
GCA pattern is approximately 9.5 minutes,
the fuel expended, approximately 740 pounds.
CCA APPROACH (Typical)
J DOWNWIND LEG
A. LamEjit" fjfnr-UP.
B. "Win# flap,*—UP*
C* Hi rut ill 1 —85% RPM ml dim mu,
D* Speed brakes—As required (o
maintain 195 knots IAS*
2 base leg
A. Land in" gear—DOWN.
B. Win# flaps—LIP*
C. Throtlle—85% RPM minimum,
D* lAS-!80 KNOTS,
Figure 9 - 7 ,
3 FINAL APPROACH
V. Wi ^ 11 aps—'T. V K VA tF F.
IL IAS-160 KNOTS,
4 GLIDE SLOPE
Use speed brakes |o maintain
I60kn»t<< IAS,
9-11
Section IX
T*0, 1F-S9H-1
NOTE:
* JFJif'rt using the flight computer , fross*
icilA bmic flight instruments.
• The time required to complete a standard
ILS approach (using the flight computer),
is approximately 6 minutes , the fuel
expended 7 approximately 880 pounds*
OUTER
MARKER
IIS APPROACH
wm rum cotAPom
(Typical)
1 OUTBOUND
A. Flight compuler selector switch—*(VOR-LOC) LEFT.
B. Altitude—As locally prescribed-
C* Altitude control switch—ON.
D. Landing gear—UP*
E. Wing flaps—IIP.
F. Throttle—BS% RPM minimum*
G. Speed brakes-As required to maintain 195 knots IAS*
2 PROCEDURE TURN
A. Begin procedure turn as locally prescribed.
B* Plight computer selector switch—FiJGHT INST,
C. Altitude control switch—OFF (prior to descent),
3 INBOUND
A* Set heading pointer to localizer heading*
B* Turn flight computer selector switch to
(VOR-LOC) RIGHT* This will cause flight
computer vertical bar to deflect*
C. When the course indicator vertical bar begins
to move off the peg, zero vertical har. This will
bring you to localizer beam.
4 intercepting glide slope
When the course indicator horizontal bar reaches
center (indicating you arc on the glide slope),
turn selector switch to APPROACH.
5 COCKPIT CHECK
A, Descend to proper altitude.
R* Landing gear—DOWN.
C, Wing flaps—TAKEOFF*
D. Throttle—85% RPM minimum* to maintain 16(1 knots IAS.
6 ON GLIDE SLOPE
Fly airplane lo center horizontal and vertical bars to
maintain position on localizer anil glide slope.
7 MISSED APPROACH
To go around in the event of a missed approach,
press the go-around button and initiate after¬
burning- By centering the bars* you will assume
a safe climbing altitude. Then follow local
missed-approach procedure*
Figure 9 - 8 .
9-12
T,0, 1F-89H-1
Section IX
MISSED-APPROACH GO-AROUND
PROCEDURE.
If a missed approach or a go-around is required, ac¬
complish the following:
L Throttles—OPEN; use afterburners for accelera¬
tion if necessary, but consideration must be given
to increased fuel consumption.
2. Speed brake lever—CLOSED.
3. Establish a takeoff or climb attitude,
4. When vertical velocity indicator and altimeter
show definite climb indication, retract gear and flaps.
5. Execute established missed-approach procedure
for the particular field.
FLIGHT COMPUTER MISSED APPROACH
WITH ILS.
Do not use flight computer missed-approach
procedure if a go-around with both after¬
burners and a clean configuration cannot be
accomplished.
If an approach has been missed on ILS and a straight¬
ahead dimbout can be made safely* the flight com¬
puter can be used to accomplish a go-around. Press¬
ing the flight computer go-around button (altitude
switch, figure 4-15) with the flight computer selector
switch at APPROACH, will displace the horizontal bar
to the optimum climbuut angle. Flying the airplane
to center the horizontal and vertical bars will then
result in a safe climbout airspeed if maximum power
is used on both engines. In the following go-around
procedure, each step should be performed without
hesitation.
1. Throtties—OPEN; use afterburners.
2. Speed brake lever—CLOSED.
3. Flight computer go-around burton—Press.
4. Landing gear lever—UP.
3. Wing flap lever—UP.
6. Fly the airplane to center the horizontal and ver¬
tical bars until desired altitude is reached. Execute
established missed-approach procedure for the par¬
ticular field.
Note
When the desired altitude is reached, the
go-around feature Is cut out by turning the
flight computer selector switch from the
APPROACH position,
INSTRUMENT LETDOWNS AND
APPROACHES ON SINGLE ENGINE.
Letdowns and approaches on single engine, either by
radar control or on the radio range, can be made
satisfactorily. Use the following procedure when mak¬
ing a single-engine GCA or ILS approach:
1. GCA downwind leg and ILS outbound—-Use 200
knots IAS, power at 88 ± 2% rpm, gear down, flaps
up, and speed brakes closed,
2, GCA base leg and ILS inbound—Use 160 to 170
knots IAS, power at 96 ± 2% rpm, gear down, flaps
up, and speed brakes closed.
3* Final approach—Use 160 to 170 knots IAS, gear
down, flaps at takeoff, power as required to maintain
desired flight path. Use 98% rpm and control rate of
descent with speed brakes.
INTRODUCTION.
The thin wings and high speeds of jet aircraft can
result in critical ice accumulation in relatively light
icing conditions in those airplanes with the anti¬
icing systems inoperative. Surface king can reduce
IAS and range of the airplane considerably. Icing oc¬
curs when the supercooled water in fog, clouds, or
rain impinges and freezes on the airplane surfaces.
Normally the heaviest icing takes place in clouds with
strong vertical currents (cumulus clouds, projections
9-13
Section IX
TO, 1 F-89H-I
above stratocumulus clouds, etc). Icing conditions
as found in stratus clouds are generally light to mod¬
erate; however, severe icing conditions may occur in
this type of cloud. Prolonged flights through moder¬
ate icing can build up as much ice as a short flight
through severe icing conditions. The most severe type
of ice formation will generally occur above ~5°C
(23°F>*
SURFACE ICING,
Surface icing normally occurs at temperatures near
0°C (32°F) on the outside air temperature gage. The
anti-icing system will keep all heated surfaces clear
of ice without noticeable loss of engine thrust. The
system will also effectively de-ice the airplane if ice
is allowed to accumulate on the wings and tail. The
purpose of the system is to prevent formation of ice;
therefore, use the system continuously whenever con¬
ditions indicate a possibility of ice. Refer to Section IV
for operating instructions on the anti-icing systems.
If the thermal anti-icing system is inoperative
and any low level flying is to be performed
under icing conditions, a higher than normal
IAS should be used. Icing will cause the stalling
speed to increase considerably; therefore ex¬
treme caution should be used, especially during
takeoff, approaches, and landings.
ENGINE ICING,
Axial flow jet engines are seriously affected by icing.
The engine air intake anti-icing is controlled by the
anti-icing switch and care must be taken to prevent
ice buildup on these surfaces since ice ingestion by the
engine can result in engine failure. Ice forms on the
inlet screens when extended and compressor inlet
guide vanes (stator) and restricts the flow of inlet air.
This causes a loss of thrust and a rapid rise in exhaust
gas temperatures. As the airflow decreases, the fuel-
air ratio increases, which in turn raises the tempera¬
ture of the gases going into the turbine. Complete
turbine failure may occur in a matter of seconds after
ice builds up in the engine air inlet. Critical ice build¬
up on inlet screens can occur in less than 1 minute
under severe conditions. With the inlet screens re¬
tracted, blocking of the air passages between the inlet
guide vanes can still occur in 4 minutes or less. The
idea that heating due to ram pressure at high speed
will prevent icing is dangerous. The heat generated at
subsonic speed is insufficient to prevent ice formation.
Engine screens should be extended after penetration
or icing has been terminated. This procedure will
minimize damage caused by large pieces of ice being
ingested into the engine.
In Below Freezing Air Temperature.
The rate of engine icing for a given atmospheric
icing intensity with outside air below freezing tem¬
perature is relatively constant up to an airspeed of
approximately 250 knots TAS. Assuming constant ic¬
ing conditions, the rate of icing increases with in¬
creasing airspeed above 250 knots. Therefore, a reduc¬
tion of airspeed to a safe minimum will reduce the
rate of engine icing in ambient temperatures of 0°C
(32°F) or below.
In Above Freezing Air Temperature*
Unlike surface icing, engine inlet icing can occur at
temperatures above freezing. Because serious inlet duct
icing can occur without the formation of ice on the
airplane external surfaces, it is necessary to understand
what causes this type of icing in order to anticipate
it, if possible, so that immediate corrective action will
be taken w r hen positive indications of engine icing ap¬
pear, When jet airplanes fly at velocities below approxi¬
mately 250 TAS at high power setting, the intake air
is sucked, instead of rammed, into the engine com¬
pressor inlet. This suction causes a decrease in air
temperature (adiabatic cooling). Under these condi¬
tions, air at a temperature above freezing may be re¬
duced to subfreezing temperature as it enters the en¬
gine. Free moisture in the air may become supercooled
and cause engine icing although no external surface
icing is evident. The maximum temperature drop
which can occur on most current engines is a drop of
approximately 5°C (9°F)- The greatest temperature
drop occurs at high rpm on the ground and decreases
with (1) decreasing engine rpm, and (2) increasing
airspeed.
Indication of Engine Icing.
The initial indication of engine icing is increased ex¬
haust gas temperature. This is usually the only indica¬
tion prior to complete engine failure. At the first sign
of engine icing turn on the engine anti-icing system
immediately. Refer to Section IV for the operation
of this system.
FLIGHT IN ICING CONDITIONS*
If a flight must be made in icing conditions, and if
either the engine or surface anti-icing system is in¬
operative, observe the following precautions:
1. Avoid known areas of icing conditions. Many
areas of probable icing conditions can be avoided by
careful flight planning and study of weather condi¬
tions.
9-14
TO. 1F-89H-1
Section IX
2. If the ambient temperature is in the range of
0°C (32°F) to 3 C C (41 °F) and water is present on the
parking ramp or runways* the inlet screens should he
retracted and the engine anti-icing system turned on
immediately upon starting the engine.
3. If possible* avoid takeoff when the temperature
is between — 10°C (14 D F) and 5°C (4l°F) if fog is
present or if the dew point is within 4°C (7°F) of
the ambient temperature. These are the conditions
under which engine icing can occur without surface
icing. When freezing rain or other icing conditions
exist at takeoff, the anti-icing switch should be
placed at TAKEOFF. The loss of thrust on takeoff
is not noticeable to the pilot. Afterburners should be
used to climb rapidly above the icing conditions.
4. If the ambient temperature is in the range of
0°C (32°F) to 5°C (4l°F), the speed of the airplane
should be maintained at 250 knots or above to lessen
the possibility of inlet duct icing due to suction ef¬
fect,
5, If icing conditions are encountered at freezing
atmospheric temperatures, immediate action should be
taken as follows: change altitude rapidly by climb or
descent in layer clouds* or vary course as appropriate
to avoid cloud formations; reduce airspeed (in freez¬
ing air) to minimize rate of tee buildup; maintain
close watch of exhaust gas temperature and reduce
engine rpm as necessary to prevent excessive exhaust
gas temperature.
INTRODUCTION.
Thunderstorms and their accompanying turbulence
should be avoided if possible. The following informa¬
tion and procedures are to be used only when flying
into a thunderstorm cannot be avoided. At altitudes
above 35,000 feet, sufficient power is nor available to
regain airspeed in level flight once it has dropped to
about 200 knots IAS. If it is noted that airspeed is
dropping below 200 knots IAS* lower the nose slightly
and maintain a descent of approximately 1000 feet per
minute until airspeed is regained. Do not use afterburn¬
ers in the storm as serious trouble could be encountered
if the airplane inadvertently went into a steep spiral.
At 30,000-foot altitude or lower* once the throttle
adjustment is made, airspeed control is not a problem
and the most serious trouble to be encountered is
severe turbulence and possible hail damage. In the
storm, the airplane should not be maneuvered inten¬
tionally, However, by observing the recommended
turbulent air penetration airspeed, a maximum ma¬
neuverability margin will be sustained at all operating
gross weights without developing prohibitive load
factors. In less severe turbulence there are no airspeed
restrictions* but maneuvering should be restricted in
proportion to the degree of turbulence.
APPROACHING THE STORM
MAXIMUM SPEED
ANY TIP TANK F
275
KNOTS-IAS
>
325 KNOTS “IAS
WITH NO TIP TANK FUEL-
9-15
Section IX
T.O. 1F-89H-1
APPROACHING THE STORM.
Prepare the airplane as follows before entering the
storm,
L Adjust power to obtain a safe and comfortable
penetration speed of 225 to 275 knots IAS, If higher
airspeeds are desired, do not exceed the following;
With ANY tip tank fuel. , . ,275 knots IAS
With NO tip tank fuel.325 knots IAS
2. Pitot heat switch—ON,
3. Anti-ice switch—FLIGHT; windshield de-ice and
defog knob—NORMAL,
4, Flight computer altitude switch—OFF,
5, At night, turn cockpit lights and thunderstorm
lights to full brightness.
COLD WEATHER PROCEDURES
INTRODUCTION.
Night flying in this airplane is the same as day flight
with the following exceptions.
NIGHT TAKEOFF.
Follow instrument takeoff procedure (with normal
reference) until a safe altitude is reached. Prior to
landing, visually check main gear down by turning
landing light on in the retracted position.
r~
CAUTION
# Taxi light does not light area near the wing
tips. Be on the alert for other airplanes, crew
chief stands, and other hazards in the taxi and
takeoff areas,
# After takeoff check altimeter, vertical veloc¬
ity indicator, and airspeed indicator, to en¬
sure positive climb and acceleration.
NIGHT LANDING.
Use the normal landing procedure.
BEFORE ENTERING THE COCKPIT. 1- Airplane covers removed.
Check to see that the following items have been ac- 2. Plugs removed from engine air intake ducts, ex-
complished: haust nozzle, and engine nacelle doors.
9-16
T,0. 1F-39H-1
Section IX
3. Visual check of bottom section of front stator
blades for evidence of ice. Engine heat on shutdown
will melt ice accumulated on previous flight; melted
ice will then re freeze in the lower section of the front
stator and rotor blades. An attempted engine start will
result in starter failure. If ice is suspected, check the
engine for freedom of rotation. If engine is not free, ex¬
ternal heat must be applied to forward engine section to
melt the ice. Start engine as soon as possible after heat
application to remove all moisture before re freezing
can occur.
4, Wing flap servo followup screw and shaft cleaned
of excessive oil and grease.
Note
Excessive oil or grease on this mechanism
can cause shaft to bind in screw and move
the servo valve spool to partially restrict hy¬
draulic flow to flap motor, causing abnor¬
mally slow movement of flaps.
longer takeoff distance requirements, in¬
creased stall speeds, poor climbout perform*
ance, and a vibration in flight that could
result in an accident,
LL Canopy jettison system, seat air, and airbrakes
serviced before each flight at temperatures below
-35°C (“31°F).
BEFORE STARTING ENGINES.
A ground power unit with two 28-volt d-c leads, each
having a capacity of 500 amperes, is required for start¬
ing engines.
L Pilot's seat—-Adjust as desired. At temperatures
below — 35°C (—31°F), heat must be applied to the
seat mechanism before the seat can be adjusted.
2, Hydraulic handpump handle—Install in pump*
In flight, the radar observer may not be able to reach
the handle in its stowed position because of his heavy
arctic clothing.
3, Hydraulic supplemental pump—Check,
5* All ice removed from fuel tank vents, static air
sources, and pitot tubes*
6* Ice and snow removed from nose wheels to pre¬
vent shimmy*
7, Fuel filters and draincocks checked for freedom
from ice and heated, if necessary, to drain condensate.
8, Oil tanks preheated, if temperature is —45°C
{ —49°F) or lower, to reduce starter loads and assure
proper lubrication. However, cold engine starts can be
made if operations warrant,
9, Shock struts checked for proper inflation, and
dirt and ice removed.
r
CAUTION
Ice should nor be chipped away because the
airplane may be damaged. Check that water
resulting from ice removal does not re freeze
on airplane surfaces, especially on control
surface hinge lines.
10. All snow and ice accumulations removed from the
wings, fuselage, and tail prior to flight.
Snow and ice that accumulate on the airplane
on the ground seriously affect the airplane's
flight performance and alter handling char¬
acteristics, These accumulations result in
Note
Under some conditions of extreme subzero
temperatures, difficulty in maintaining nor¬
mal hydraulic pressure during supplemental
pump check may occur. Operation of pump
for from 3 to 5 minutes should provide nor¬
mal pump operation.
4* In extremely low temperatures, below — 40°C
40°F), apply heat to the back side of the land¬
ing gear handle mechanism to clear any ice from the
selector valve cable and prevent possible cable slip¬
page.
STARTING ENGINES.
Follow normal starting procedure outlined in Section
II. When the engine reaches 1G% rpm,open the throttle
halfway and return to IDLE. This additional move¬
ment of the throttle loosens any connections that have
become stiff, bur does not alter the fuel flow. Oil pres¬
sure may be high after starting cold engines. This is
not dangerous unless the pressure remains high. Delay
takeoff until the pressure drops to normal.
CAUTION
When ambient temperature is 0°C (32°F) or
below, have hot air from a portable heater
blown into the engine air intake ducts and
exhaust nozzles for 10 to 15 minutes. This
procedure prevents the starter-generator unit
from being damaged due to ice seizure of
the compressor rotor.
9-17
Section IX
T.O. 1 F-89H-1
GROUND TESTS,
Because of increased air density at low ambient tem¬
peratures, thrust developed at all engine speeds is
greater than normal. For ground tests at low tempera¬
tures use the following procedures;
1. Generator—Check output and make all checks
requiring electrical power before having external
powe r d i scon nected.
2. Cabin heat, windshield heat, and canopy defog—
As required.
To prevent cracking of the windshield glass,
keep windshield heat switch at NORMAL
for at least 1 minute before turning to EMER.
Never keep windshield heat switch at EMER
longer than necessary,
3, Flight controls—Check operation. At tempera¬
tures below ^35°C ( 31°F), operate flight controls
three or four times during engine runup until flight
controls operate freely and easily.
At very low temperatures, hydraulic packing
may fail and cause hydraulic leaks. Have
ground personnel check flight control mech¬
anism access doors for signs of excessive leak¬
age.
4. Wing flaps—Check operation.
5. Speed brakes—Check operation and cycle several
times to assure free movement.
6. Instruments—Check operation. Flight instruments
require approximately 2 minutes for warmup.
In cold weather, make sure that all instru¬
ments have warmed up sufficiently to en¬
sure normal operation. Check for sluggish
instruments during taxiing.
TAXIING INSTRUCTIONS.
When taxiing in cold weather, observe the following
precautions:
!. Avoid taxiing in deep snow because taxiing and
steering are very difficult, and the brakes may freeze.
2. Taxi very slowly on icy or wet surfaces; the air¬
plane is difficult to control during a skid.
3. Maintain directional control with nose wheel
steering.
Under freezing conditions, use caution when
actuating nose wheel steering on taxiing out
of parking area or after landing. Nose wheel
may be frozen in deflected position.
Note
The airplane has a strong tendency to
weathervane when taxiing on ice; however,
the steerable nose w j heel wall greatly facilitate
directional control.
To preserve the battery, use only essential
electrical equipment while taxiing at low
engine speeds.
4. When taxiing behind another airplane on icy'
taxi ways, allow enough distance between airplanes to
stop safely and to prevent icing of the airplane sur¬
faces by melted snow and ice in the jet blast of the
preceding airplane.
5. When fine powder snow is on the taxi way, the
preceding airplane's jet blast will cause a large blind¬
ing cloud of flying snow; the distance between air¬
planes must be increased for visibility.
6. Minimize taxi time to conserve fuel and to re¬
duce amount of fog generated by jet engines.
7. At very low temperatures, operate flight controls
frequently.
BEFORE TAKEOFF,
When the taxiway is covered with ice, a full power
check may not be possible before takeoff because the
airplane may slip on the ice. In this case, the power
check can be made at the start of the takeoff run by
opening the throttles rapidly and turning on the after¬
burners. If afterburners do not ignite on both en¬
gines, discontinue takeoff. Very low temperatures do
not appreciably affect rudder and elevator operation.
However, at temperatures below 35 °C { 31°F), the
ailerons become stiff and should be cycled several
times before takeoff to ensure easy movement.
1, Rocket heater switch—ON if mission requires use
of rockets,
2. Anti-icing system—ON if necessary.
9-18
T.O. TF-89H-1
Section IX
During takeoff the anti-icing switch should
not be used in the FLIGHT position unless the
runway will allow' a 20 to 25 percent longer run
than required for a normal takeoff. This is due
to the reduction of engine thrust caused by
anti-icing hot air being bled {at a very high
rate) from the llth stage of the engine com¬
pressors whenever rhe anti-icing system is
used w'itfa the switch placed in the FLIGHT
position.
3* Fuel filter de-ice switch—Hold at each position
for approximately 10 seconds to remove any accumu¬
lation of ice.
TAKEOFF,
At rhe start of rhe takeoff run, advance the throttles
rapidly and turn on afterburners to make power check.
If afterburner on either engine does not ignite, do not
take off. After a rakeoff from a snow or slush covered
field, operate the landing gear, wing flaps, and speed
brakes several times to remove slush and snow that
might cause these units to freeze in the streamlined
positions.
Do not exceed landing gear and flap struc¬
tural airspeed limitations.
Arctic flight tests have shown that light frost accu¬
mulations have no effect on takeoff and disappear at
250 knots IAS. At very low temperatures, do not apply
brakes after takeoff to stop the wheels spinning be¬
cause the brakes may freeze in the braked position.
Depending on the weight of snow and ice
accumulated, takeoff distances and climb-out
performance can be seriously affected. The
roughness and distribution of the ice and
snow j could vary stall speeds and characteris¬
tics to an extremely dangerous degree. Loss
of an engine shortly after takeoff is a serious
enough problem without the added, and
avoidable, hazard of snow and ice on the
wings. In view of the unpredictable and un¬
safe effects of such a practice, the ice and
snow r must be removed before flight is at¬
tempted.
DURING FLIGHT.
Flight characteristics are unchanged by arctic condi¬
tions except for aileron stiffness at temperatures be¬
low ^35°C (“3l°F). The ailerons should be oper¬
ated periodically throughout the flight if these tem¬
peratures are encountered. If only the left hydraulic
system is operating, the rudder should also be oper¬
ated periodically. Turn on de-icing and anti-icing
systems as needed. Check all instruments since some
instruments may be unreliable at 1ow t temperatures.
Before penetration, fuel filter de-icing should be used
for 10 seconds in each fuel system to de-ice the filters
and engine fuel controls.
Note
Engine fuel control icing will cause the fuel
flowmeter to fluctuate. This indicates that
flameout of an engine may be imminent,
APPROACH TO PATTERN.
At temperatures be!ow r — 35°C ( — 31°F), operate the
ailerons several times before entering the pattern to
ensure smooth and easy operation. Follow' normal pat¬
tern and approach procedures, but allow for longer
approach than normal because high thrust at low
temperature results in a flatter glide. Wing flap ex¬
tension requires 2 seconds longer than normal, and
retraction requires 7 seconds longer than normal at
—65°F. Speed brake operation requires a maximum
of 1.5 seconds additional time to open or close at
— 65°F. Normal landing gear extension and retraction
requires 2 seconds longer at 65 °F; however, emergency
extension requires 25 seconds longer.
Note
• When making GCA approaches during arc¬
tic operations, decrease power settings about
3 percent because of increased thrust at Iow r
temperatures.
• The windshield and canopy defrost systems
should be operated at the highest temperature
possible (consistent wdth the pilot’s comfort)
during high-altitude flight in order to pro¬
vide Sufficient preheating of the transparent
surfaces to preclude the formation of frost or
fog during descent.
• On initial approach use alcohol on each en¬
gine for IQ seconds.
9-T9
Section IX
TO. 1F-89H-1
LANDING*
a ■ iPiAft I
CAUTION
Operation of anti-icing system during landing
affords protection against icing conditions
but causes loss of thrust. If a go-around is
necessary, the anti-icing switch may remain
in the FLIGHT position only if two engines
w'ith maximum thrust and afterburning are
available.
For minimum landing roll on wet or icy runways,
both the wing flaps and speed brakes should be fully
extended during landing roil* and the right engine
should be shut down immediately after three wheel
contact. Open the speed brakes after main gear touches
down and leave extended until after turning off run¬
way. The aerodynamic drag of the wing flaps and
speed brakes partially offsets the decreased braking
efficiency experienced when landing on wet and icy
runways and the thrust eliminated by shutting down
the idling right engine will aid in reducing the landing
roll. Apply brakes carefully and intermittently after
touchdown. If the airplane has snow-and-ice tires, ap¬
ply brakes carefully and intermittently after touchdown
to prevent tread from filling and glazing over. Glazing
reduces braking effectiveness on icy runways, and land¬
ing ground roll distances may be increased as much as
100 percent more than the distances shown in the Land¬
ing Distance Chart (see Appendix),
BEFORE LEAVING AIRPLANE*
Check that ground personnel perform the following:
1. Service airplane as soon as possible.
2. Remove dirt and ice from shock struts,
3. Clear snow and ice from nose wheels.
4. Service canopy jettison system and airbrake bot¬
tle if temperature is below 35 C ( 31°F) and the
airplane is to he used for another flight.
5. Check flight control access doors for signs of
excessive hydraulic leakage.
6. Install plugs in engine air intake ducts, exhaust
nozzles, and engine nacelle doors.
7. Cover pitot tubes and all static air sources.
8. Check fuel pumps, filters, and draincocks for ice
and drain condensate within 30 minutes after stopping
engines.
9. Bleed and recharge engine screen pneumatic
system.
10, Install covers on wings, empennage, and canopy.
11, Remove battery and store in a heated room if
layover of several days is anticipated, or if tempera¬
ture is below — 29°C (^20°F),
INTRODUCTION*
Takeoff and landing rolls are longer In hot weather
because of the lower air density which also lengthens
takeoff rolls by decreasing engine performance. Added
precaution should be taken to protect rubber and plastic
parts of the airplane from damage by excessive heat,
BEFORE ENTERING THE AIRPLANE*
Check tires for blisters, abrasions, proper inflation, and
excessive wear. Be sure external ground cooler is dis¬
connected,
TAKEOFF*
Anticipate a longer takeoff distance than normal. Re¬
fer to Appendix I, figure A-6 for takeoff distances.
AFTER TAKEOFF—CLIMB*
Be sure to maintain specified climbing airspeed, cor¬
recting maximum rates of climb as required by the
effects of high temperatures on rates of climb encoun¬
tered under hot weather flight conditions. Refer to
Climb Chart in Appendix.
LANDING*
Anticipate longer landing distances and use minimum
wheel braking to prevent overheating of brakes. Refer
to Appendix I, figure A-29 for applicable landing
distance charts.
BEFORE LEAVING AIRPLANE*
Be sure canopy is protected from direct rays of the
sun.
9-20
T.O, 1F-89H-1
Section IX
,
1 DESERT PROCEDURES
.- -• ^ , HF-36B ^*1
BEFORE TAKEOFF.
Do not run engines during a dust or sand storm unless
absolutely necessary. Before engine runup, position
the airplane so it will not receive dust from, or blow
dust on, other airplanes.
TAKEOFF.
Avoid takeoff in blowing dust or sand.
INTRODUCTION.
When operating under desert conditions, the normal
hot weather procedure is followed. In addition, pre¬
cautions must be taken to prevent external abrasion
of the airplane surfaces and to keep sand and dust
from entering the airplane systems.
BEFORE ENTERING THE COCKPIT.
L Check exposed shock struts and actuating cylin¬
ders for dust and sand. Have them cleaned if neces¬
sary.
2. Check all air intakes for sand and dust.
3. Check wheel brake disks for excessive abrasion.
BEFORE LEAVING AIRPLANE.
Close and seal the canopy during dust or sand storms,
and check that ground personnel perform the follow¬
ing:
L Cover canopy to prevent sand abrasion.
2. Cover all air intakes and ducts as soon as pos¬
sible after landing.
r
9*21
1,0. 1F-89H-1
Appendix I
PCRFORMANCl
TABLE OF CONTENTS
Page
INTRODUCTION . A-1
CORRECTION TABLES. A-4
PERFORMANCE CHARTS. A-4
TYPICAL MISSION . A-21
Airspeed Position Correction. A-25
Compressibility Correction to Calibrated
Airspeed ..... . A-26
Temperature Correction for Compressibility. . A-27
Density Altitude Chart ..A-28
Takeoff and Landing Crosswind Chart . . . , . A-29
Takeoff Distance Maximum Power. A-30
Critical Field Length A-36
Refusal Speeds .. , , . . A-37
Velocity During Takeoff Ground Run
Maximum Power ...... .. , * * A-38
Minimum Distance Climb ..* *.A-40
Best Climb Performance (Range) Maximum
Power . , , ,..,. A-4T
Nautical Miles Per TOOO Pounds Fuel Sea Level A-65
Mission Profile Basic Plus Pylons ..A-86
Intercept Profile Basic Plus Pylons ..A-89
Optimum Return Profile Basic Plus Pylons . . . A-92
Maximum Endurance Basic Plus Pylons .... A-95
Optimum Maximum Endurance Profile
Basic Plus Pylons , , . ... . . ..A-98
Descents , .,.. ,.. *.A-101
Landing Distance . . * , *.A-103
Landing Speeds .. . . A-107
Combat Allowance Chart—Maximum Power .A* 108
INTRODUCTION.
The flight performance charts in this section provide
the pilot with flight planning data and airspeed and
ambient temperature correction data. Two types of
performance charts are included: profile-type charts for
maximum range, endurance, and continuous power
operation, and graphical charts for takeoff, climb,
nautical miles per 1000 pounds of fuel, descents, and
landings,
PROFILE CHARTS.
T he profile-type charts are a supplement to the graphi¬
cal data and help flight planning by reducing the
computations that must be made. These charts are
based on the recommended climb and cruise settings
shown on the profile for the particular configuration
involved and give direct indication of the fuel and
time required to cover a given distance if the recom¬
mended settings are adhered to. For flight planning
based on settings other than those given on the profile
charts, the graphical charts should be used. Decreased
weight due to fuel consumption has been accounted for.
GRAPHICAL CHARTS.
The graphiail charts provide detailed performance data
for one- and two-engine operation. T hese charts should
be used for flight planning when performance data not
covered in the profile charts is needed. Unless otherwise
indicated, all data pertains exactly to NACA standard
ambient temperatures but may be considered approxi¬
mate for nonstandard conditions. The CAS or Mach
number tabulated for each pressure altitude should be
maintained for nonstandard temperatures regardless of
the deviations of other quantities from the given
values, except when it is necessary to use a lower CAS
value or Mach number to avoid exceeding engine limits.
TAKEOFF DISTANCE
MODEL F-89H MAXIMUM POWER ENGINE(S): (2) J35-35
DATA BASIS- FLIGHT TEST wtTH 08 WITHOUT PYLON TANKS FUEL GRADE: JP-4
DATE: 22 OCTOBER 1957 FUEL DENSITY; 6.5 LB/US GAL
Appendix 1
T,G, 1F-89H-I
Sample.
A-2
T.O. 1F-89H-1
Appendix I
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A-3
Sampfe,
REMARKS 1 (JSE 30 DEGREE Ft APS 3. USE 1Q&% RPM WITH AFTERBURNING UNLESS LIMITED BY
2' DISTANCE SHOWN WILL BE OBTAINED WHEN TAKEOFF IS IN ACCORDANCE WITH MAXIMUM TAILPIPE TEMPERATURE
SPECIFIED NORMAL PROCEDURE. ON DRY hard SURFACE RUNWAY ‘ ENGINE AIR INLET SCREENS EXTENDED,
Appendix )
TO, 1F-89H-1
CORRECTION TABLES*
AIRSPEED CORRECTIONS.
Assuming zero instrument error, the pilot's airspeed
indicator reads correct indicated airspeed (IAS). Cor¬
rections must be applied to IAS to determine calibrated
airspeed (CAS), equivalent airspeed (BAS), and true
airspeed (TAS). The algebraic sum of the installation
correction and IAS equals CAS. The CAS value minus
the compressibility correction equals EAS. EAS divided
by the square root of the relative air density (V^)
equals TAS. Relative air density is equal to the ratio
of the free airstream ambient density at altitude to
standard sea level density. Wind velocity added vec¬
tor ially to TAS equals ground speed (GS). Corrections
to be applied to convert IAS to CAS are tabulated in
the Airspeed Position Correction Table (figure A-l),
These corrections are given for values of IAS and
pressure altitude for the operating range of the clean
configuration; corrections for flap settings and gross
weights are also shown. Landing gear position does not
affect airspeed readings. Values for converting CAS to
EAS are shown in the Compressibility Correction to
Calibrated Airspeed Table {figure A-2) which covers
the operating CAS and pressure altitude range of the
airplane. Values of the reciprocal of the square root of
the relative air density (1 -4- Vo?), used for determining
TAS, are obtained from the Density Altitude Chart
(figure A-4). The airspeed indicator in the radar
observer’s cockpit indicates approximate TAS; there¬
fore, only the wind correction need be applied to
determine ground speed.
AMBIENT TEMPERATURE CORRECTIONS*
A compressibility correction must be applied to the
temperature gage reading to obtain true ambient tem¬
perature. The correction is shown as a function of CAS
and pressure altitude in the Temperature Correction for
Compressibility Table (figure A-3).
USE OF THE CORRECTION TABLES.
Assume the following instrument readings:
1. Altimeter 35,000 ft
2. Airspeed indicator 284 kn
3. Free air temperature gage — I9°C
The correct airplane speed and ambient temperature
are:
4. IAS (zero instrument error)
284 kn
5. Installation correction
+ 5 kn
6. CAS
289 kn
7, CompressIbility correctioo
— 18 kn
8, EAS
271 kn
9. Free air temperature gage reading
- 19°C
10. Temperature correction for
compressibility error.
— 25°C
11. Correct ambient temperature
—44°C
At 35,000-foot pressure altitude and — 44°C, the re¬
ciprocal of the square root of the relative air density
(1 Vo:) from figure A-4 is 1.85. Therefore* TAS is
271 X 1.85 = 501 knots.
TAKEOFF AND LANDING CROSSWIND CHART*
A Takeoff and Landing Cross wind Chart (figure A-5)
enables the pilot to convert crosswind to a component
headwind down the takeoff or landing runway. The
component headwind is used to accurately determine
takeoff ground run and landing ground roll. The
Takeoff and Landing Crosswind Chart is also used to
determine if takeoff or landing is recommended under
crosswind conditions at the predicted minimum nose-
wheel liftoff and touchdown speeds.
Use of Takeoff and Landing Crosswind Chart*
When the wind direction and velocity and runway
heading are known, the component headwind down
the takeoff runway can be determined from the Take¬
off and Landing Crosswind Chart. With a wind from
330 degrees at 20 knots velocity and using runway 27,
the chart is entered at (330 degrees — 270 degrees) 60-
degree angle and 20-knot wind velocity, Reading to
the left, the component headwind down the takeoff
runway is found to be 10 knots.
To determine if takeoff is recommended under the
above conditions, proceed vertically from the inter¬
section of runway wind angle and crosswind lines to
the predicted takeoff airspeed of 134 knots. Takeoff
is found to be recommended,
PERFORMANCE CHARTS*
TAKEOFF DISTANCE CHARTS.
The Takeoff Distance Charts (figure A-6) show takeoff
distances (ground roll and total distance to clear a
50-foot obstacle) as a function of gross weight, pressure
altitude, wind velocity, and ambient temperature for a
dry, hard-surface runway. Gross weight, wind velocity,
and ambient temperature are always known factors;
the pressure altitude of the field can be determined
by setting the altimeter to 29-92 (sea level standard
day pressure in inches of mercury). The charts show
data for two-engine takeoffs with maximum or military
power, using the norma] procedure given in Section IL
If an engine fails during military power takeoff, after¬
burning on the operating engine should be started im¬
mediately or the takeoff discontinued. Military power
data may be used to estimate adequate field length if
afterburners fail during takeoff.
Note
Takeoff with military power will result in
a fuel saving of only 250 pounds. This fuel
saving will result in an increased range of
only 25 nautical miles. The slight increase in
range must be weighed against the additional
risks involved in military power takeoffs.
T.O. 1F-S9H-1
Appendix i
Single-engine maximum power takeoff dam is also
included to determine the required takeoff distance
when power on one engine is lost during takeoff (see
Section III). If the takeoff technique used is different
from that specified in Section II, the distances will
differ from those shown in the charts. A deviation of
5 percent from the airspeeds in Section II will result
in a distance deviation of 10 percent or more.
Use of Takeoff Distance Charts,
The Takeoff Distance Sample Chart shows a maximum
power takeoff at an ambient air temperature of 15°C,
pressure altitude of 2000 feet, gross weight of 40,000
pounds and a 20-knot headwind. This results in a
ground roil of 2500 feet and a total distance of 4000
feet to clear a 50-foot obstacle.
CRITICAL FIELD LENGTH CHART,
The Critical Field Length Chart (figure A-7), in con¬
junction with the Refusal Speed Chart (figure A-8),
can be used to determine a course of action if an engine
fails at any point during the takeoff ground run for
any combination of critical field and runway lengths.
For example, comparison of the critical field length
with the runway length available indicates the follow¬
ing takeoff limitations:
Runway Length Greater Than Critical Field Length,
L At engine failure speeds below refusal speed:
If the runway is longer than necessary for one-engine
takeoff, the pilot has the option of either taking off or
stopping. If the runway is shorter than necessary for
one-engine takeoff, pilot must stop.
2. At engine failure speeds above refusal speed, pilot
must take off, as stopping within the limits of the
rumvay is impossible.
Critical Field Length Greater Than Runway Length.
1. At engine failure speeds below refusal speed, pilot
must stop, as takeoff within the limits of the runway
is impossible.
2. At engine failure speeds above refusal speed, the
pilot must take off with remaining engine.
Use of Critical Field Length Chart,
The Critical Field Length Sample Chart shows a maxi¬
mum power takeoff with ambient air temperature of
15°C, a pressure altitude of 2000 feet, a gross weight
of 40,000 pounds, and a 20-knot headwind. These con¬
ditions indicate a critical field length of 4600 feet.
According to the Takeoff Distance Chart (figure A-6)
for one-engine takeoff, the runway length required for
one-engine takeoff is 6800 feet. If the available runway
length is 6000 feet, the refusal speed is found to be
109 knots IAS. Thus, the available runway length is
greater than the critical field length but shorter than
necessary for one-engine takeoff. UndeF these condi¬
tions, if the speed at the point of engine failure is less
than 109 knots IAS, the pilot should stop the airplane
rather than attempt a one-engine takeoff; if the speed
at the point of engine failure is greater than 109 knots
IAS, the pilot should take off, as stopping within the
limits of the runway would not be possible,
REFUSAL SPEED CHART.
The Refusal Speed Sample Chart shows a maximum
mum speed at which engine failure permits stopping
at the end of the runway. It is based on normal takeoff
procedure and a dry, hard-surface runway.
Use of Refusal Speed Chart,
The Refusal Speed Sample Chart show's a maximum
power takeoff at a gross weight of 46,000 pounds, a
pressure altitude of 2000 feet with an ambient air
temperature of 59° F, and a 7000-foot runway. The
resulting refusal speed is 114 knots.
VELOCITY DURING TAKEOFF GROUND RUN
CHARTS,
The Velocity During Takeoff Ground Run Charts
(figure A-9) are based on normal operating procedures
as specified in Section II and show the relationship
bet^veen indicated airspeed and distance traveled dur¬
ing takeoff ground run on a dry, hard-surface runway.
These charts are useful for checking takeoff accelera¬
tion by reference to a go-no-go marker located a known
distance from the end of the runway. This is deter¬
mined by subtracting distance remaining at go-no-go
marker from runway available. On an odd length run¬
way, one half of the odd figure over exact thousands
of feet must be added to the distances shown on the
markers to determine the actual distance remaining.
This distance is used to enter acceleration curves (fig¬
ure A-9) to determine go-no-go speed. Since accelera¬
tion check marker is two markers short of go-no-go
marker, the acceleration check speed is determined at
a distance 2000 feet less than go-no-go distance.
Use of Velocity During Takeoff Ground Run Charts,
Enter the chart at the applicable gross weight of the
airplane. Read over to the base line, then proceed
vertically downward to the required takeoff ground run
distance as determined from the Takeoff Distance
Charts (figure A-6). From this point trace a curve
parallel to the guide lines until it intersects the distance
being used as a checkpoint. This point shows the
velocity which should be attained at that distance.
In the Velocity During Takeoff Ground Run sample
chart, the takeoff gross weight is 43,000 pounds, the
required takeoff distance at maximum power is 3500
feet, and the distance from the start of the takeoff run
ro the acceleration checkpoint is 1500 feet. The result¬
ing velocity at the checkpoint is 84 knots IAS, and the
takeoff velocity is 136 knots IAS,
A-5
Appendix I
T.O. 1F-89H-?
MODEL* F-89H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
CRITICAL HUD LIHCTH
WITH OH WITHOUT PYLON TANKS
MAXIMUM POWER
ENGIN£{$); (2) J 35-35
FUEL GRADE: JP-4
FUEL DENSITY: 63 LB/US GAL
REMARKS.
1 ALL VALUES SHOWN ON CHART ARE BASED ON DRY HARD SURFACE RUNWAY 30-DECREE FLAPS. AND SPEED
BRAKES INOPERATJVE.
2 THREE SECONDS ALLOWED FOR PILOT RECOGNITION OF ENGINE FAILURE; AT THE END OF THE THREE
SECONDS, THROTTLES ARE CUT AND BRAKES APPLIED
3. ENGINE INLET SCREENS EXTENDED.
SompJe.
SAMPLE CHART
DO NOT USE FOR
FLIGHT PLANNING
A-6
T.O. 1F-89H-1
Appendix i
MODEL F-89H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
mmt spms
maximum power
WITH OR WITHOUT PYLON TANKS
ENGINE(S): (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY:6.5 LB/US GAL
REMARKS:
1. ABOVE VALUES ARE BASED ON DRY HARD SURFACE RUNWAY, USING SPECIFIED NORMAL
TAKEOFF PROCEDURE UP TO POINT OF ENGINE FAILURE AND OPERATION IN
ACCORDANCE WITH SECTION III AFTER ENGINE FAILURE-
2. ENGINE AIR INLET SCREENS EXTENDED.
'T
him*
Sample.
A-7
Appendix J
T.O. IF-89H-1
MODEL: F-S9H
DATA BA$fS: FLIGHT TEST
DATE: 22 OCTOBER 1957
REMARKS' INDICATED AfRSPEED KNOTS
1 ^? CITlES SHOWN WIU BE OBTAINED WHEN TAKEOFF 15 tN ACCORDANCE W[TH SPECIFIED NORMAL PROCEDURE,
2 ENGINE AIR INLET SCREENS EXTENDED
H-JD3
Sample.
A-8
moc/ry during takeoff ground run
MAXIMUM POWER
WITH 08 WITHOUT PYLON TANKS
ENGINE(S): (2) J35-35
FUEL GRADE; JP-4
FUEL DENSITY: 6.5 LB/US GAL
GROUND ROLL DISTANCE-1000 FT
T*0, TF-89H-I
Appendix I
MINIMUM DISTANCE CLIMB CHART.
Depending on gross weight and thmst, minimum dis¬
tance climb (maximum angle of climb) at low altitudes
may he obtained at the applicable airspeeds shown in
figure A-10,
USE OF MINIMUM DISTANCE CLIMB CHARTS,
Enter the applicable configuration chart at the in¬
tended gross weight and read up to the proper inter¬
secting thrust line. From the point of intersection of
gross weight and thrust lines, follow to the left and
read minimum distance climb airspeed from the left
side of the chart. For a climb following takeoff, initial
climb weight is the takeoff gross weight minus the
906-pound takeoff fuel allowance,
BEST CLIMB CHARTS.
The Best Climb Charts (figures A-ll through A-19)
show climb performance in terms of fuel, time, air
distance, rate of climb, and climb CAS necessary to
attain this performance. Data is given for climbing
with two engines at maximum, military, and normal
power, and with one engine at maximum and military
power. The fuel, time, and air distance values shown
include the effects of kinetic energy change and weight
reduction during climb, but do not include any allow¬
ance for start, takeoff, or acceleration. Time and dis¬
tance are plotted against gross weight with guide lines
to show the reduction in gross weight during climb
due to fuel consumption. In most cases, three charts are
provided for each configuration and power setting:
these include two Best Climb Performance Charts (one
plotted against distance, the other plotted against time):
and one Best Climb Speed Chart (showing rate of climb
and best climb CAS)*
Use of Best Climb Charts,
To obtain the desired data from the Best Climb Charts,
enter the proper climb chart at the gross weight and
altitude at start of climb and note the time (or distance)
and fuel used at this point* From this initial point,
trace a curve parallel to the guide lines until it inter¬
sects the desired altitude at end of climb. Note the
time (or distance) and fuel used at this intersection.
The difference between the initial and final time is
the time required to climb. The difference between the
initial and final values for distance and for fuel used
gives, respectively, the distance traveled and fuel used
in climb. Since time, distance, and fuel used in climb
are zero at sea level, these values may be read directly
for climbs starting at sea level. It must be kept in
mind, however, that for a climb following takeoff, the
initial climb weight is the takeoff gross weight minus
the 906-pound takeoff fuel allowance. The appropriate
sample shows the fuel used and time to climb from
10,000 feet to 35,000 feet using military power with
pylon tanks and a gross weight of 41,000 pounds at
start of climb. Rate of climb and best climb CAS may
be obtained directly from the Best Climb Speed Charts.
TAKEOFF DATA CARDS.
A Takeoff Data Card (see Abbreviated Checklist, Sec¬
tion II) is to be completed before each flight. The
purpose of the takeoff data card is to familiarize the
pilot with emergency procedures to be followed in the
event of engine failure or other emergencies which
may occur on takeoff. Critical field length, refusal
speed, acceleration checkpoint speed, and the other in¬
formation required on the takeoff data card may be
found in the Appendix charts*
Use of Takeoff Data Cards.
Sample Problem, Assuming that takeoff flaps are
used and that the center of gravity is within limits, the
following conditions are given preparatory to com¬
pleting the Takeoff Data Card that follows:
TAKEOFF DATA
Gross Weight 40,000 Lb Pressure Altitude 2000 Ft
Runway Length 8000 Ft Headwind 20 Kn
Temperature 59*F Surface (Dry, Wet, Icy)
Takeoff Distance. . .Normal 2700 Ft 50-ft Obstacle 3500 Ft
Takeoff Distance. .. 1 Engine 7500 Ft 50-ft Obstacle 12,300 Ft
Critical Field Length 4600 Ft Refusal Speed 131 Kn
TAKEOFF (Maximum Power)
Acceleration Check 75 Knots IAS at 1000 Fr
Nose Wheel Liftoff Speed. ....****.,,.*, 124 Kn
Takeoff Speed ...... 129 Kn
Initial Climb Speed (To Dear 50-foot Obstacle) ..... 141 Kn
Decision Factors;
1* Critical field length is less (greater or less) than
runway length.
2. If engine failure occurs at a speed below maximum
refusal speed, you should abort the takeoff *
3* If engine failure should occur at a speed in excess
of refusal speed, you should proceed with maximum
power on operating engine and use engine failure dur¬
ing takeoff procedure .
NAUTICAL MILES PER 1000 POUNDS FUEL CHART,
Cruise data throughout the normal speed range may be
obtained from the Nautical Miles Per 1000 Pounds
Fuel Charts (figures A-20, A-21, and A-22). Each chart
includes specific range (nautical miles per 1000 pounds),
fuel flow (pounds per hour), and power settings
(% rpm), as well as curves of maximum endurance and
recommended long-range cruise speeds for zero wind.
Specific range is plotted against Mach number, with
subscales of calibrated airspeed (CAS) and true airspeed
(TAS),
Use of Nautical Miles Per T 000 Pounds Fuel Charts.
To obtain the cruising range for a given amount of
fuel, use the following steps:
L Select the proper chart for the airplane configu¬
ration and altitude.
A-9
Appendix l
T.O. 1F-89H-1
MODEL F-89H
DATA BASIS; FLIGHT TEST
DATE: 22 OCTOBER 1957
BEST cum PERFORMANCE (TIME)
ENGINE(S): (2) J 35-35
FUEL GRADE; JP-4
FUEL DENSITY; 6.5 LB/US GAL
MILITARY POWER
PYLON TANK CONFIGURATION
1. FUEL CONSUMPTION INCREASED 5 PERCENT FOR SERVICE VARIATION.
7 . SUBTRACT 906 POUNDS FROM AVAILABLE FUEL TO ALLOW FOR WARMUP,
TAXI, AND TAKEOFFj ENTER CHART AT TAKEOFF GROSS WETGHT LESS
906 POUNDS.
3. ENGINE AIR INLET SCREENS RETRACTED.
OPTIMUM CRUISE ALTITUDE NORMAL RATED POWER,
SAMPLE CHART
DO NOT USE FOR
FLIGHT PLANNING
Sample.
A-10
T.O. 1F-89H-1
Appendix 1
NAUTICAL MtUS PTR WOO POUNDS FUU
MODEL; F-89H
30,000 FEET
BASIC CONFIGURATION PLUS PYLONS
ENGINEfS): (2) J35-35
DATA BASIS: FLIGHT TEST
DATE- 22 OCTOBER 1957
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB/US GAL
1 FUEL CONSUMPTION INCREASED 3 PERCENT TO ALLOW FOR SERVICE VARIATION.
2 ENGINE AIR INLET SCREENS RETRACTED.
3. MAINTAIN CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE.
SAMPLE CHART
DO NOT USE FOR
FLIGHT PLANNING
H
Sample,
A-Tl
Appendix I
T.O. 1F-89H-1
2. Determine the average weight of the airplane for
the amount of fuel being considered.
3, Enter the graph at this average weight and the de¬
sired Mach number, or desired power setting (% rpm),
to obtain specific range {nautical miles per 1000 pounds
of fuel).
4, Multiply the specific range by the amount of fuel
(pounds -T- 1000) to obtain cruising range.
5. Interpolate the approximate fuel flow and power
setting (% rpm) at the Mach number and average
weight.
Sample Problem. Determine the range obtainable
from 6000 pounds of fuel at an altitude of 30,000 feet
and long-range cruise speed. The long-range cruise
speed is the higher of two speeds for a given altitude
and gross weight where 99% of the maximum range is
obtainable.
With an initial airplane weight of 40,000 pounds and
basic configuration plus pylons:
1. Select the proper chart for the airplane
configuration and altitude.
2. Find the average weight
, 6000
(40,000 - —- )
37,000 lb
3- Enter the chart at the intersection of the
zero wind cruise line and 37,000 pounds
gross weight and read:
Specific range 100.4 n mi per 1000 lb fuel
Mach number 0.681
RPM 90%
Fuel flow 4000 lb per hr
The range is then found
(100,4 + 1000 X 6000)
602 n mi
MISSION PROFILE CHARTS,
The Mission Profile Charts (figure A-23) show the
relationship of time, fuel, distance, and altitude to
maximum range for no-w r ind conditions. This relation¬
ship is based on a mission sequence of takeoff, military
power climb, and long-range cruise. The fuel curves
include a 906-pound allowance for start, taxi, and
takeoff, the fuel used in climbing to each altitude, and
the fuel required for long-range cruise. The time lines
include the time required for climbing to cruise alti¬
tude, but do nor include the time for start, taxi, or
takeoff. The line labeled Initial Climb Path shows the
distance traveled during the military power climb from
sea level to cruising altitude, using the climb speed
schedule tabulated at the left of the chart. The con¬
tinuation of the initial climb path is the cruise-climb
path based on a constant Mach number. The approxi¬
mate best cruise-climb altitude can be obtained by
climbing at the recommended military power schedule
until the rate of climb is 500 feet per minute, then
leveling off and setting up the recommended power
setting and Mach number. The airplane will automati¬
cally seek the cruise-climb altitude for its particular
gross weight. The initial throttle setting should be
maintained throughout the remainder of cruise-climb.
For cruise at a constant altitude, the recommended
Mach number should be set up at the intersection of
the climb path and the cruise altitude. As the flight
progresses, the power setting must be decreased grad¬
ually to maintain the recommended Mach number as
fuel is consumed. As an aid to preflight planning, a
line of best range for constant-altitude flight appears
on the chart. This curve is not a flight path, but a plot
for best cruise altitude against distance. For distances
greater than those covered by the curve, cruise-climb
procedure for maximum range should be used, A cruise
table gives recommended Mach numbers and approxi¬
mate operating conditions for both cruise-climb pro¬
cedure and cruise at constant altitude.
5. Average speed is Mach no. X speed of
sound (0.681 X 589) 401 kn
6. Time in cruise may be determined by
nautical miles -5- knots
(602 -5- 401) 1.51 hr or 1 hr 31 min
When wind conditions are encountered, the air nautical
miles per 1000 pounds of fuel read from the chart may
be converted to ground nautical miles per 1000 pounds
of fuel as follows:
ground N MI _ air N Ml ^ V ground
1000 pounds 1000 pounds V air
where
V air = airplane true airspeed
V ground = airplane true ground speed —
V air ± V wind
Use of Mission Profile Charts.
The charts may be entered with one or more of the
four range factors of time, fuel, distance, and altitude.
By entering the chart with the known factors, the
others may readily be determined for a no-wind con¬
dition. To determine wind effect upon time, fuel, and
distance, compute the average true airspeed (distance -r
time, no wind) and apply wind to TAS to obtain
ground speed (GS). Then compute the time with wind
(distance -r GS). Reenter the profile at the cruising
altitude and the computed time with wind to determine
the fuel required with wind.
Sample Problem T * Using the Mission Profile sample
chart, find the fuel required, time, necessary speed, and
power setting to cruise 250 nautical miles at 20,000 feet
against a headwind of 40 knots with no external load.
A-12
TO. 1F-S9H-T
Appendix 1
L Enter at 250 n mi and 20,000 ft to
obtain fuel required (no wind) 4800 lb
2. Time (no wind) 40 min
(0,67 hr)
3. Calculate average TAS (250 + 0.67) 375 kn
4. Apply wind to obtain GS (375 — 40) 335 kn
5. Calculated time with 40-kn headwind
(250 h- 335 ) 45 min
(0.75 hr)
6 . Reenter at cruise altitude at the time
with wind. Fuel required with wind 5200 lb
Sample Problem 1* Using the Intercept Profile sam¬
ple chart, find the fuel required, time, necessary speed
and power setting to cruise 200 nautical miles at 25,000
feet against a head wind of 40 knots with no external
load.
1 . Enter at 200 n mi and 25,000 ft to
obtain fuel required (no wind) 5400 lb
2 , Time (no wind) 25 min
(0.42 hr)
3. Calculate average TAS (200 -H 0,42) 475 kn
4, Apply wind to obtain GS
(475 - 40) 435 kn
7. Tabular cruise speed 0.6l Mach
no.
8 . Tabular cruise power setting 87% rpm
(approx)
Note
If this flight had been made at 26,500 feet
cruising altitude (reference, the line of best
range at 250 nautical miles), the time and fuel
required would have been less.
Sample Problem 2 , Determine the maximum distance
flyable with no external load, 10,000 pounds of fuel,
and a 60 -knot headwind.
Enter at 1 0,000 lb of fuel and obtain
maximum air distance at
cruise-climb (no wind) 835 n mi
Time (no wind)
2 hr 4 min
(2,07 hr)
Calculated average TAS
(835 -h 2.07)
403 kn
Apply wind ro obtain GS (403
60 ) 343 kn
Calculate distance with wind
(2.07 X 343)
710 n mi
Tabular cruise-climb speed
0.70 Mach no.
INTERCEPT PROFILE CHARTS.
The Intercept Profile Charts (figure A*24) present the
fuel required to fly a given distance in a minimum of
time, consistent with reasonable range capabilities.
These charts are based on maximum pow’er climb and
military power cruise; they are similar to the Mission
Profile Charts and are used in the same manner.
Notice, however, that use of the Intercept Profiles
should be restricted to flights that require a minimum
of time, whereas the Mission Profile Charts are used
for maximum range flights.
5, Calculated time with 40-kn wind
(200 435) 28 min
(0.46 hr)
6 . Reenter at cruise altitude at the time
with wind, fuel required with wind 6000 ib
7. Tabular cruise speed .81 Mach no.
8, Tabular cruise power setting 100% rpm
Sample Problem 2, Determine the maximum distance
flyable with no external load and 10,000 pounds of fuel
and a 60-knot headwind,
I, Enter at 10,000 Ib of fuel and obtain
maximum air distance at cruise-climb
(no wind)
690 n mi
2. Time (no wind)
1 hr 32 min
(1,53 hr)
3. Calculated average TAS (690 -5- 1.53)
450 kn
4. Apply wind to obtain GS (450 60 )
390 kn
5. Calculate distance with wind
(1.53 X 390)
600 n mi
6 , Tabular cruise-climb speed
*77 Mach no.
OPTIMUM RETURN PROFILE CHARTS,
The Optimum Return Profile Charts (figure A-25)
show the minimum fuel required for maximum dis¬
tance (no wind) based on an optimum flight path from
any point within the range of the airplane configura¬
tion. The flight path required is indicated by the
different shaded areas and the notes relative to them.
The fuel curves are based on a military power climb to,
and recommended cruise at, the optimum altitude. The
military power climb speed schedule and recommended
cruise settings are tabulated on each chart. No reserve
for loiter, descent, or landing has been included. The
time shown at the optimum altitude is cruise time only;
it does not include the time required for climb to
optimum altitude or any allowance for loiter, descent,
or landing.
A-13
Appendix I
T.O. 1F-89H-1
I
Sample.
A-14
TO* 1F-89H-1
Appendix V
/
)
Sample.
A-15
Appendix 1
T.O. 1F-89H-1
Use of Optimum Return Profile Charts.
MAXIMUM ENDURANCE CHARTS.
The chart may be entered at the initial altitude with
either the fuel on board (to determine the distance
available) or with the distance to be flown (to deter¬
mine the fuel required)* The shaded area in which the
initial point falls establishes the necessary procedure,
as stated in the note relative to the area, to obtain
maximum range. The time required to fly the distance
is the time at cruise altitude (obtained from the profile)
plus the time required to climb, if necessary (obtained
from the Military Power Climb Chart for the applica¬
ble configuration)* The effect of wind must be applied
to obtain the actual fuel and time to fly the distance*
A close approximation can be obtained by considering
the head or tailwind for the time it requires to com¬
plete the flight (neglecting the difference in wind at
the lower altitudes since comparatively little time is
spent during the climb phase).
Sample Problem. From the Optimum Return Profile
sample chart, determine the fuel and time required to
return to a base 800 nautical miles away. The airplane,
carrying pylon tanks, is at 20,000 feet with 10,000
pounds of fuel on board (grossweight— 4 1,957 pounds).
A 60-knot headwind is assumed.
1. Enter profile at 800 n mi and 20,000 ft
to establish starting point. Fuel
required (no wind) 8000 lb
2. In this area, note that a climb is
required and a cruise-climb procedure
followed*
3- Following the climb path guide lines,
the initial cruise altitude is 31,000 ft
4, Cruise time (no wind) 1 hr 50 min
5. From the military power chart for
pylon tank configuration, time
to climb 13 min
6* Total time (no wind; H 4”+"5 ,> ) 2 hr 3 min
7. Average TAS (distance -r- total time) 390 kn
8. Average ground speed
(TAS — headwind) 330 kn
9. Total time with headwind
(distance + average ground speed) 2 hr 25 min
10* Cruise time with wind ( ,f 9”—2 hr 12 min
11* Using the cruise time "10” on the
profile, back track down the climb
path from the line of best range to
20,000 ft to obtain fuel required
with wind 9650 lb
12. Fuel remaining over base at altitude
(10,000 - 9650) 3501b
13* Use the flightpath originally
determined for no wind*
The Maximum Endurance Charts (figure A-26) show
the maximum time available with the fuel on board
when loitering at a constant altitude. The recommended
calibrated airspeed and the approximate operating
conditions are tabulated on each chart.
Use of Maximum Endurance Charts.
To determine the time available for a given amount of
fuel, enter the chart at the amount of fuel on board at
the start of loiter and the flight altitude and note the
initial time. Reenter the chart at the amount of fuel
on board at the end of the endurance flight (initial
fuel on board less fuel to be used) and read the final
time. The difference between the initial and final time
is the time available to loiter at constant altitude. To
obtain the fuel required to loiter a given time, enter
the chart at the amount of fuel on board at the start of
loiter and flight altitude and note the initial time.
Reenter the chart at time of end of loiter (initial time
less time to loiter) and read final fuel on board. The
difference between the initial and final fuel on board
is the fuel required to loiter.
Sample Problem. From the Maximum Endurance
sample chart, determine the fuel required to loiter at
30,000 feet with no external load for 45 minutes* The
fuel on board at start of loiter is 6000 pounds (gross
weight—37,677 pounds)*
1. Initial time at 6000 lb and 30,000 ft
1 hr 56 min
2, Final time (1:56 — 0:45)
I hr 11 min
3* Fuel on board at end of loiter
(1:11 at 30,000 ft)
3550 ib
4. Fuel required to loiter
(6000 lb - 3550 lb)
2450 lb
5. Recommended loiter CAS
195 CAS
OPTIMUM MAXIMUM ENDURANCE PROFILE
CHARTS.
The Optimum Maximum Endurance Profile Charts
(figure A-27) give the maximum time in the air with
the fuel remaining, based on an optimum flight path
from any starting altitude. The flight path required is
indicated by the different shaded areas and the notes
relative to them. Time and fuel lines shown are based
on a normal power climb (military power climb in the
case of one-engine operation) to best endurance alti¬
tude, loiter at that altitude, and a maximum range
descent to sea level (no reserve for landing). The climb
speed schedule is tabulated at the left of the chart; the
loiter speed schedule is tabulated below the chart.
Use of Optimum Maximum Endurance Profile
Charts.
The chart may be entered at the initial altitude with
either the fuel remaining (to determine the time avail¬
able) or the time desired (to determine the fuel
A-l 6
T.O. 1 F-S0H-T
Appendix I
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Appendix l
T.Q* 1F-89H-1
Sample,
A-18
T.O. 1F-89H-I
Appendix 1
requirement).The shaded area in which the initial point
falls establishes the flight path to be used, as stated in
the note relative to the area.
Sample Problem- Using the Optimum Maximum En¬
durance Profile Sample Chart, determine the time
available and the necessary flight path for maximum
endurance aloft in the pylon tank configuration with
6000 pounds of fuel remaining at 20,000 feet.
L Enter profile at 20,000 ft and 6000 lb
of fuel remaining to establish starting
point. Total time available 1 hr 55 min
2. In this area note that a climb is
required,
3. Follow the climb path guide lines for
the best endurance altitude
4. Descent time from 27,600 ft to
sea level
5. Elapsed time from start of climb to
start of descent ("1” —"4")
If a reserve of 1000 lb of fuel is desired
for landing, enter the profile at 6000 lb
of fuel and follow the climb path guide
line to the best endurance altitude
6. Subtract endurance time due to the
1000-lb fuel reserve (at altitude for
best endurance) 18 rnin
7. Descent time from 27,600 ft to
sea level
8. Elapsed time from start of climb to
sea level C6”-K7”)
of the field can be determined by setting the altimeter
to 29-92 (sea level standard day pressure in inches of
mercury). The chart for two-engine operation shows
data for landing using the normal procedure given in
Section II. The chart for one-engine operation is based
on inoperative speed brakes and flaps. If the landing
technique used differs from that specified, the landing
distances w r iil vary from those shown on the charts,
A 5-percent variation in speed causes approximately a
10-percent variation in distances; insufficient wheel
braking may increase ground roll by 50 percent.
Use of Landing Distance Charts.
The Landing Distance sample chart shows a landing
with two engines operating at an ambient air tempera¬
ture of 15°C and a pressure altitude of 2000 feet with a
gross weight of 32,000 pounds and a 20-knot headwind.
These conditions require a ground roll of 2250 feet and
a total distance of 3250 feet from a 50-foot obstacle
clearance to end of ground roll,
LANDING IMMEDIATELY AFTER TAKEOFF
DATA CARD.
A Landing Immediately after Takeoff Data Card is
to be completed before each takeoff. The purpose of
the landing immediately after takeoff data card is to
familiarize the pilot with emergency procedures to be
followed if loss of an engine or other emergencies
necessitate landing immediately after takeoff. Infor¬
mation necessary to complete the normal landing and
single-engine landing sections may be found in the
Appendix charts.
27,600 ft
23 min
1 hr 32 min
27,600 ft
I hr 14 min
23 min
1 hr 37 min
DESCENT CHARTS.
The Descent Charts (figure A-28) show descent per¬
formance for one and two engines operating in terms
of fuel, time, air distance, and rate of descent for the
gross weight range of the airplane denoted by the
shaded areas. Charts are shown for no external load and
maximum external stores configuration. Interpolation
must be used for intermediate configurations and
gross weights. The type of tip pod has negligible effect
on descent. Three types of descents are shown: recom¬
mended descent with speed brakes closed (based on
0,70 Mach number), recommended descent with speed
brakes open (based on 0.70 Mach number), and maxi¬
mum range descent (based on approximately 200 knots
IAS). All three types of descent are based on idle
power. These charts may be used for descending from
one altitude to another by taking the incremental
values between the initial and final altitudes,
LANDING DISTANCE CHARTS.
The Landing Distance Charts (figure A-29) show land¬
ing distances (ground roll and total distance to clear a
50-foot obstacle) for a dry, hard-surface runway as a
function of gross weight, pressure altitude, wind ve¬
locity, and ambienr temperature. The pressure altitude
Use of Landing Immediately After Takeoff
Data Card.
Sample Problem, The following conditions are given
as a basis for completing the normal and single-engine
landing sections of the sample Landing Immediately
after Takeoff Data Card,
I
| LANDING IMMEDIATELY AFTER TAKEOFF DATA |
I Maximum Emergency Landing Weight ................ 38.000
(Takeoff Weight Less Jeccisonable Items)
I Engine 2 Engine
I Final Approach Speed Ill Kn 153 Kn
1 Touchdown Speed ... 139 Kn 122 Kn j
j Ground Roll Distance 3900 Ft 2900 Ft I
Total Distance (To Clear 50doot Obstacle)- 7000 Ft 4000 Ft |
I______—--1
LANDING DATA CARD.
A Landing Data Card (see Abbreviated Checklist, Sec¬
tion II) is to be completed before each flight. The
purpose of the landing data card is to familiarize the
pilot with emergency procedures to be followed if loss
of an engine or other emergencies occur during land¬
ing, The information required by the normal landing
and single-engine landing sections of the landing data
card may be found in the Appendix charts.
A-I9
Appendix E
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T.O. 1F-89H-1
Appendix I
Use of Landing Data Cards*
Sample Problem- The following conditions are given
as a basis for completing the normal landing and
single-engine landing sections of the sample Landing
Data Card*
LANDING DATA
Landing Gross Weight . * *.38+000 Lb
Kunway Length 8000 Ft Headwind 20 Kn
Temperature 5P fl C Pressure Altitude 2000 Ft
Surface: (Dr)', Wet* Icy)
LANDING
l Engine 2 Engine
Final Approach Speed . +■+..+.. 162 Kn 140 Kn [
Touchdown Speed . . . .... 128 Kn 113 Kn
I Landing Distance . . ^. ....++..+ 3200 Ft 2450 Fc I
[ Landing Distance (To Clear 50-foot Obsrade) + . 6000 Ft 3500 Ft I
__I
LANDING SPEEDS CHART*
The Landing Speeds Chart (figure A-30) presents the
recommended indicated airspeeds for final approach,
50-foot obstacle clearance, touchdown, and nose wheel
down. The chart may be read for applicable landing
gross weights and for flap settings of 0 degrees, 30 de¬
grees, and 50 degrees.
COMBAT ALLOWANCE CHARTS*
The Combat Allowance Charts (figure A-31) show the
relationship between time and fuel with changes in
altitude for two-engine operation at maximum, mili¬
tary* and normal power. Combat time or fuel may be
determined from this chart for a given power setting*
Use of Combat Allowance Charts*
Enter the chart at the combat altitude and the fuel
quantity to be used for combat to obtain the time
available. Enter at the altitude and time available for
combat to obtain the fuel required.
TYPICAL MISSION.
This sample problem combines the use of the charts in
this section to plan a typical mission,
FLIGHT PLAN DATA*
A combat mission is to be flown carrying pylon tanks
on the inbound leg, the tanks to be dropped at the
beginning of combat. Prepare a flight plan based on
the following data:
L Distance to combat area 400 nmi
2, Assigned altitude:
Inbound to combat (cruise-climb) 28,000 ft
and above
Outbound from combat 33,000 ft
(cruise-climb) and above
3. Combat at 40,000 ft
(Maximum power) 10 min
4 . Weather (assume standard day
temperature throughout) CAVU
Winds aloft inbound (28,000 ft
and above) 40-kn HW
Winds aloft outbound (33,000 ft
and above) 50-kn TW
Field elevation 2000 ft
5 * Airplane gross we i gh t:
Operating minimum (includes crew
of two, oil, trapped fuel, pylons, and
miscellaneous equipment) 30,155 lb
Forty-two 2.75" FFAR rockets 760 ib
Six GAR-1 missiles 762 lb
Two 300-gallon pylon tanks 280 lb
Maximum usable fuel—internal and
external (2369 gallons) 15,398 Ib
Total gross weight 47,355 1b
TAKEOFF*
Obtain takeoff distance from the maximum power
takeoff distance chart, figure A-6+ (Standard day tem¬
perature at 2000 feet is 11°C) Assume 20-knot head¬
wind.
L Ground roll distance (47,355 lb) 4900 ft
2+ Total takeoff distance over 50-ft obstacle 5800 ft
3+ Takeoff speed (IAS) 144 kn
INBOUND LEG*
Cruise*
The inbound leg may be determined directly from the
Mission Profile Chart for pylon tanks carried through¬
out, figure A-23, since at a distance of 400 nautical
miles at the cruise-climb altitude some fuel remains in
the ranks. The profile includes a 906-pound fuel allow¬
ance for start, taxi, and takeoff, as well as the fuel,
time, and range required for climb to and cruise at the
cruise-climb altitude.
Distance
400 n mi
Fuel required (no wind) from profile 6750 lb
Time (no wind) from profile
I hr 2 min
Average TAS ("T*-7***3")
387 kn
Ground speed ("4"— 40 kn)
347 kn
Time with wind CT”-^“5")
1 hr 9 min
Fuel required (with wind) from
profile
7400 lb
Cruise speed (cruise-climb altitude)
.68 Mach no.
Cruise power setting
94% rpm
(approx)
Military power climb speed schedule
(see figure
A-16, Sheet 2
of 3 Sheets)
Gross weight at end of cruise
(47,355 Ib —"7")
39,955 lb
A-21
DATA BASIS: FLIGHT TEST FUEL GRADE: JP-4
DATE: 22 OCTOBER 19S7 FUEL DENSITY: 6.5 LB/US GAL
Appendix f
T.O, 1F-89H-1
A-22
LANDING DISTANCE TO CLEAN SOFT-OBSTACLE
WITH OR WITHOUT PYLON TANKS *
T*0» 1F-89H-1
Appendix l
Sample
A-23
Appendix 1
10, TF-89H-1
Climb to Combat Altitude.
Maximum power climb to combat altitude (40,000 ft).
1. Distance traveled in climb
2, Gross weight at start of climb
35 n mi
<29,590 ft)
3. Gross weight at end of climb to
39,955 lb
40,000 ft
38,900 lb
4. Fuel used to climb (39,955 — 38,900)
1,055 lb
5. Time to climb
5 min
6. Maximum power climb schedule
7, Drop pylon tanks at end of climb
(gross weight at begining of combat
(See figure
A-16, Sheet 1
of 3 Sheets)
is "3”- 280)
COMBAT,
38,620 lb
From the combat allowance chart (figure A*31, Sheet 1
of 3 Sheets), obtain the fuel required for combat at
40,000 feet.
1. Combat—maximum power (10 min)
2. Gross weight at end of combat
38,620 — 1800 lb (combat fuel)
~ 762 lb (six GAR-1 missiles)
- 760 lb (forty-two2,75" FFAR
1800 lb
rockets)
35,298 lb
Assume zero distance traveled during combat. Deter*
mine the fuel remaining at end of combat.
3. Takeoff, climb, and cruise
7400 lb
4. CHmb to combat altitude
1055 lb
5. Combat
1800 lb
6. Total fuel used
10,255 lb
7. Fuel remaining (15,398 — 10,255)
51431b
OUTBOUND LEG.
Cruise-Climb.
At the end of combat the airplane is 435 nautical miles
400 ■+■ 35) from the base at an altitude of 40,000 feet.
Enter the Optimum Return Profile Chart (figure A-25,
Sheet 1 of 3 Sheets) for basic configuration + pylons
at the distance from the base and determine the fuel
required and reserve with the existing tailwind. Note
that optimum altitude for start of return at the distance
is 34,400 feet; therefore, a recommended descent (with
speed brakes open) is made from 40,000 feet to 33,300
feet (time, distance, and fuel consumed are negligible).
I. Distance
435 n mi
2. Fuel required (no wind)
36001b
3* Initial cruise altitude
33,300 ft
4. Total time (no wind)
1 hr 4 min
5. Average TAS ( f T’^ r '4 M )
408 kn
6. Average ground speed ("5”+ 50 kn) 458 kn
7. Total time with wind ("l”-!-'^”)
57 min
8. Fuel required (with wind)
3300 lb
9* Cruise speed
10. Power setting (See figure A-20,
OJOMachno.
Sheet 8 of 9 Sheets)
92% rpm
(approx)
11. Reserve fuel over base (5143 —"8")
at 35,600-ft altitude
1843 lb
Descent.
Obtain the fuel required to descend to
base from the
Descent Chart (figure A-28, Sheet 1 of 2 Sheets).
1- Recommended descent, speed brakes
open from 35,600 ft
50 lb
2. Time to descend
1 min
3. Desce n t speed, us i ng id 1 e power
and speed brakes open
4. Fuel reserve for loiter and landing
0.70 Mach no.
(1843 — 50)
1793 lb
5, Airplane gross weight for landing
31,948 lb
Landing,
Obtain the landing distance from the Landing Distance
Chart (figure A-29, Sheet 1 of 4 Sheets). Use 2000-foot
altitude, 11°C and 204tnot headwind.
L Ground roll distance
2920 ft
2. Total distance over 50*ft obstacle
4080 ft
3, Approach speed (IAS)
150 kn
4. 50-ft obstacle speed (IAS)
127 kn
5. Touchdown speed (IAS)
The sum of all the time required gives
119 kn
the time from takeoff to landing
2 hr 22 min
A-24
T.O. 1 F-89H“1
Appendix
nmspeeo position connection
MODEL: F- 89 H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
ENGINE(S); (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB/US GAL
REMARKS:
1. ADD CORRECTION TO CORRECTED INSTRUMENT READING ilASi TO OBTAIN CALIBRATED AIRSPEED.
2 . GEAR UP OR DOWN.
H3T1
Figure A-L
COMPRESSIBUITy CORRECTION to calibrated airspeed
Appendix I
TO. 1F-89H-1
SION* - S A — V
A-2 6
CALIBRATED AIRSPEED -V c -KNOTS
TEMPERATURE CORRECTION FOR COMPRESSIBILITY
TO, 1F-B9B-1
Appendix I
A-27
CALIBRATED AIRSPEED —V c - KNOTS
DENSITY ALTITUDE-tOOO FT
Appendix I
T.O. 1F-89H-1
_j_
TEMPERATURE-“C
DtHSITy AtT/TVDC CHARI
—- 1,06
— 1 -04
2.30
2.20
2.10
2.00
1,26
1.22
no
— 1 os
\ ,90
1.86
1,82
1 78
1.74
1,70
1.66
1.62
1.53
1.54
1.50
1.46
1.42
1.3B
1.34
1.30
1 -28
1.24
1.20
1.16
M2
— 1.36
— 1,32
— 1.75
—1.88"
— 1.B4
— 1,80
1,76
1.72
— 1 68
— 1,64
— 1,60
1.56
1,52
— ! .40
— 1.44
— 1.40
Figure 4-4.
T.O. 1F-89H-1
Appendix I
* - IAS KNOTS
Appendix !
TO. 1F-S9H-1
? O £
H* i
w <
% :
Jt oe ?
65*5
§ 2 :
H, 1 C
§l s
i s ?
o
o
£
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5” J-
< <
O Q
Figure A-6 /Sheet 1 of 6),
A-30
TAKEOFF DISTANCE TO CLEAR SOFT-OBSTACLE
T.O. 1F-89H-1
Appendix 1
Figure A-6 (Sheet 2 of 6).
A-31
REMARKS; ) USE GO-DEGREE Ft. A PS 3- USE 1&&3S RPM WtIH AFTERBURNING UNLESS LIMITED BY
2 DISTANCE SHOWN WILL 8E OBTAINED WHEN TAKEOFF t$ IN ACCORDANCE WITH MAXIMUM TAlLP?FE TEMPERATURE
SPECIFIED NORMAL PROCEDURE, ON DRY HARD-SURFACE RUNWAY. A ENGINE AIR INLET SCREENS EXTENDED
Appendix E
T.O. 1F-89H-1
Figure A-6 fSheef 3 of 6).
A-32
T,0* 1F-89H-1
Appendix I
<
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an
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i
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be;
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LO
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Figure A-6 fSheef 4 of 6).
A-33
REMARKS: ] USE 30-DEGREE FLAPS. 3 USE 100% RPM UNLESS LIMITED BY MAXIMUM TAILPIPE TEMPERATURE
2 DISTANCE SHOWN WSLL BE OBTAINED WHEN TAKEOFF IS 4. IF ONE ENGINE FAILS DURING TAKEOFF IMMEDIATELY START AFTERBURNER
IN ACCORDANCE WITH SPECIFIED normal PROCEDURE ON ON operating ENGINE OR discontinue TAKEOFF.
DRY H AftD'SURFACE RUNWAY S, ENGINE AIR INLET SCREENS EXTENDED.
Figure A-6 (Sheet 5 of 6).
A-34
TO. 1F-89H-I
Appendix L
<
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3
j&g
a c „
1U i «
k lU 7
< ^
2 <£
“ I “
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:*§
^ £
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3 it IU
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« < o
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flf ^
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r: U ~7
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O » u.
Q a O “
So o
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£“^2
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□t J t <
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w 5 1 Q .
Cl O ~T Un
6;0£^
C JJ
IU IU Ul S£
55**;*
<
£
Figure A-6 (Sheet 6 of 6).
A-35
Appendix ]
T.O. 1F-89H-T
MODEL: F-89H
DATA BASIS; FLIGHT TEST
□ATE: 22 OCTOBER 1957
CffltCAl FIELD LENGTH
WITH OR WITHOUT PYLON TANKS
MAXIMUM POWER ENGINES): (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB/US GAL
'' BRAKES L |NOPEH AWE ^ CHART A * E 8ASH> ° N 0RY HARD ' SURfA CE RUNWAY, 30-DEGREE FLAPS, AND SPEED
2. THREE SECONDS ALLOWED FOR PILOT RECOGNITION OF ENGINE FAILURE; AT THE END OF THE THREE
SECONDS, THROTTLES ARE CUT AND BRAKES APPLIED.
3. ENGINE AIR INLET SCREENS EXTENDED
H3iS
Figure A~7.
A-36
T.O. 1F-89H-1
Appendix 1
MODEL; F-89H
DATA BASIS; FLIGHT TEST
DATE: 22 OCTOBER 1957
REFUSAL SPEEDS
MAXIMUM POWER
WITH OR WITHOUT PYLON TANKS
ENGINE(S); (2} J 35-35
FUEL GRADE; JP~4
FUEL DENSITY: 6,5 LB/US GAL
REMARKS;
1. ABOVE VALUES ARE BASED ON DRY HARD-SURFACE RUNWAY, USING SPECIFIED NORMAL
TAKEOFF PROCEDURE UP TO POINT OF ENGINE FAILURE AND OPERATION IN
ACCORDANCE WITH SECTION 111 AFTER ENGINE FAILURE,
2. ENGINE AIR INLET SCREENS EXTENDED,
Figure A-8*
A-37
Appendix I
T.O. 1F-89H-1
Figure A-9 (Sheel I of 2).
GROUND ROLL DISTANCE-1000 FT
T.O. 1F-89H-1
Appendix I
mocny during tarboff ground run
MODEL: F-89H
DATA BASES: FLIGHT TEST
DATE n OCTOBER T957
MILITARY POWER
WITH OR WITHOUT PYLON TANKS
ENGINE'S}; (2) J 35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB/US GAL
1 VELOCITIES SHOW hi will be obtained when takeoff is in accordance with specified normal procedure
2 ENGINE AIR INLET SCREENS EXTENDED
it
Figure A-9 (Sheet 2 of 2 !j.
A-39
GROUND ROLL DISTANCE-1000 FT
IAS-KN0T5 IAS—KNOTS
Appendix I
T.O. 1F-89H-1
MODEL: F-89H
Data BAstS: flight test
DATE: 72 OCTOBER 1957
MINIMUM DISTANCE CUMB
SEA LEVEL TO 10,000 FT
ENGlNECSh (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB/US GAL
34 35 36 37 38 39 40 41 42 43 44 45 46
GROSS WEIGHT-1000 LB
TWO 300-GALLON PYLON TANKS
GROSS WEIGHT-1000 LB
A-40
Figure A- 10.
T.O. 1F-89H-1
Appendix i
BEST CUMB PERFORMANCE (RANCC)
MAXIMUM POWER
MODEL: F-B9H BASIC CONFIGURATION PLUS PYLONS
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
ENGINE(S): (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 IB/US GAL
GROSS WEIGHT — 1000 LB
1. FUEL CONSUMPTION INCREASED 5 PERCENT FOR SERVICE VARIATION
2. SUBTRACT 9Q6 POUNDS FROM AVAILABLE FUEL TO ALLOW FOR WARMUP r
TAXI, AND TAKEOFF; ENTER CHART AT TAKEOFF GROSS WEIGHT LESS
?06 POUNDS,
3. ENGINE AIR INLET SCREENS RETRACTED.
* OPTIMUM CRUISE ALTITUDE - NORMAL RATED POWER,
H3?2
Figure A-ll (Sheet ? of 3),
A-4I
Appendix \
T.O. 1F-89H-1
MODEL F-S9H
DATA BASIS: FLIGHT TEST
DATE- 22 OCTOBER 1957
BEST CUm PERFORMANCE (RANGE)
MILITARY POWER
BASIC CONFIGURATION PLUS PYLONS
ENGINE(S): (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB/US GAL
GROSS WEIGHT - 3 000 LB
REMARKS:
1. FUEL CONSUMPTION INCREASED S PERCENT TO ALLOW FOR SERVICE VARIATION,
2. SUBTRACT 906 POUNDS FROM AVAILABLE FUEL TO ALLOW FOR WARMUP,
TAXI, AND TAKEOFF, ENTER CHART AT TAKEOFF GROSS WEIGHT LESS
906 POUNDS,
3. ENGINE AIR INLET SCREENS RETRACTED.
* OPTIMUM CRUISE ALTITUDE - NORMAL RATED POWER,
hjt3
A-42
Figure A-11 (Sheef 2 of 3).
T.O. 1F-89H-1
Appendix !
MODEL F-89H
DATA BASIS: FLIGHT TEST
DATE: n OCTOBER 1957
BIST CUMS PCRfORMAHCt (KMClj
NORMAL POWER
BASIC CONFIGURATION PLUS PYLONS
ENGINE(S): (2) J35-35
FUEL GRADE; JP-4
FUEL DENSITY: 6,5 LB/US GAL
GROSS WEIGHT - }Q0Q 16
REMARKS!
*■ FUE1 CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION
2 SUBTRACT 906 POUNDS FROM AVAILABLE FUEL TO ALLOW FOR WARMUP.
TAX!. AND TAKEOFF; ENTER CHART AT TAKEOFF GROSS WEtGHT LESS
906 POUNDS,
3 ENGINE AIR INLET SCREENS RETRACTED,
* OPTIMUM CRUJSE ALTITUDE NORMAL RATED POWER,
urn
Figure A-1J (Sheet 3 of 3).
A-43
Appendix !
TO* 1F-B9H-1
MODEL E-89H
DATA BASIS; FLIGHT TEST
DATE: 71 OCTOBER 1957
BEST cum PERFORMANCE (TME)
MAXIMUM POWER
BASIC CONFIGURATION PLUS PYLONS
ENGINES); (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB/US GAL
GROSS WEIGHT - 1000 LB
REMARKS:
1. FUEL CONSUMPTION INCREASED 5 PERCENT FOR SERVICE VARIATION.
2. SUBTRACT 906 POUNDS FROM AVAILABLE FUEL TO ALLOW FOR WARMUP,
TAXI, AND TAKEOFF; ENTER CHART AT TAKEOFF GROSS WEIGHT LESS
906 POUNDS.
3. ENGINE AIR INLET SCREENS RETRACTED.
* OPTIMUM CRUISE ALTITUDE - NORMAL RATED POWER.
H32S
Figure A-12 (Sheef T of 3J.
A-44
TO, 1F-89H-1
Appendix f
MODEL: F-89H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER ?957
BEST COMB PERFORMANCE (me)
MILITARY POWER
BASIC CONFIGURATION PLUS PYLONS
ENGINES): (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6,5 LB/US GAL
3ROSS WEIGHT - 1000 LB
REMARKS:
1. FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION,
2. SUBTRACT 906 POUNDS FROM AVAILABLE FUEL TO ALLOW FOR WARMUP,
TAXI, AND TAKEOFF; ENTER CHART AT TAKEOFF GROSS WEIGHT LESS
906 POUNDS
3. ENGINE AIR INLET SCREENS RETRACTED
* OPTIMUM CRUISE ALTITUDE NORMAL RATED POWER,
H3J4
Figure A-12 (Sheet 2 of 3}.
A-45
Appendix I
T.O. 1F-89H-1
MODEL: F-89H
DATA BASIS: FLIGHT TEST
DATE; 22 OCTOBER 1957
BEST CUMB PERFORMANCE (TtMB)
NORMAL POWER
BASIC CONFIGURATION PLUS PYLONS
ENGINES):® J35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB/US GAL
REMARKS:
1 FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION,
2 SUBTRACT 906 POUNDS FROM AVAILABLE FUEL TO ALLOW FOR WARMUP,
TAXI, AND TAKEOFF; ENTER CHART AT TAKEOFF GROSS WEIGHT LESS
906 POUNDS,
3„ ENGINE AIR INLET SCREENS RETRACTED.
* OPTIMUM CRUISE ALTITUDE - NORMAL RATED POWER.
K37?
Figure A-12 (Sheet 3 of 3),
A-46
Appendix
MODEL: F-891
DATA BASIS:
DATE 72 OC
RATE OF CLIMB - TOGO FT/MIN
I* CLIMB AT CAS SHOWN REGARDLESS OF AMBIENT TEMPERATUHE-
2. ENGINE AIR INLET SCREENS RETRACTED.
BEST CLIMB SPEED-KNOTS CAS
Figure A-13 (Sheet 2 of 3J*
T.O. 1F-89H-1
Appendix 1
MODEL: F-89H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
BEST cum SPEED
ENGlNEfS); (2) J35-35
FUEL GfiSADE: JP-4
FUEL DENSITY: 6,5 LB/US GAL
NORMAL POWER
BASIC CONFIGURATION PLUS PYLONS
REMARKS:
BEST CUMB SPEED - KNOTS CAS
1, CLIMB AT CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE.
2, ENGINE AIR INLET SCREENS RETRACTED.
Figure A-13 fSheef 3 of Z).
A-49
Appendix !
T.O. 1F-89H-1
BEST CUMB PERFORMANCE (RANGE}
MAXIMUM POWER
MODEL: F-89H PYLON TANK CONFIGURATION
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
ENG1NHSJ: (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB/US GAL
REMARKS:
GROSS WEIGHT - 1000 LB
1. FUEL CONSUMPTION INCREASED 5 PERCENT FOR SERVICE VARIATION,
2. SUBTRACT 906 POUNDS FROM AVAILABLE FUEL TO ALLOW FOR WARMUP,
TAXI. AND TAKEOFF, ENTER CHART AT TAKEOFF GROSS WEIGHT LESS
906 POUNDS.
3. ENGINE AIR INLET SCREENS RETRACTED.
^OPTIMUM CHUJSE ALTITUDE - NORMAL RATED POWER.
H33T
Figure 4-T4 fSheef T of 3^,
A-50
T*0* 1F-89H-1
Appendix 1
MODEL F-89H
DATA BASIS: FLIGHT TEST
DATE» 22 OCTOBER T957
BEST cum PERFORMANCE (RANGE)
MILITARY POWER
PYLON TANK CONFIGURATION
ENG!NE[S): (2) J35-35
FUEL GRADEJP-4
FUEL DENSITY: 6.5 LB/US GAL
GROSS WEIGHT - 1000 LB
REMARKS:
1. FUEL CONSUMPTION INCREASED 5 PERCENT FOR SERVICE VARIATION,
2 . SUBTRACT 906 POUNDS FROM AVAILABLE FUEL TO ALLOW FOR WARMUP,
TAXI, AND TAKEOFF- ENTER CHART AT TAKEOFF GROSS WEIGHT LESS
906 POUNDS.
3. ENGINE AIR INLET SCREENS RETRACTED.
* OPTIMUM CRUISE ALTITUDE - NORMAL RATED POWER r
nan
Figure A-1 4 (Sheet 2 of 3J,
A-5T
Appendix I
T.O. 1F-89H-1
MODEL F-OTH
DATA BASIS: FtIGHT TEST
DATE: 22 OCTOBER 1957
BEST CUMB PERFORMANCE (tuuke)
NORMAL POWER
PYLON TANK CONFIGURATION
ENGINE(S): (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY: 63 LB/US GAL
4B 44 44 42 40 38 36 34 32 30
GROSS WEIGHT — 1000 LB
REMARKS:
1. FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION,
2- SUBTRACT 906 POUNDS FROM AVAILABLE FUEL TO AUOW FOR WARMUP,
taxi, and takeoff enter chart at takeoff gross weight less
906 POUNDS.
3, ENGINE AIR INLET SCREENS RETRACTED.
* OPTIMUM CRUISE ALTITUDE — NORMAL RATED POWER.
H333
Figure A-14 (Sheet 3 of 3).
A-52
T.O. 1F-89H-1
Appendix I
MODEL F-89H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
best cam pmoumHct amej
MAXIMUM POWER
PYLON TANK CONFIGURATION
ENGlNEfS): (2) J35-35
FUEL GRADE; JP-4
FUEL DENSITY: 6.5 LB/US GAL
GROSS WEIGHT - 1000 LB
REMARKS;
1. FUEL CONSUMPTION INCREASED 5 PERCENT FOR SERVICE VARIATION,
2 . SUBTRACT 906 POUNDS FROM AVAILABLE FUEL TO ALLOW FOR WARMUP,
TAX]. AND TAKEOFF; ENTER CHART AT TAKEOFF GROSS WEIGHT LESS
906 POUNDS,
3. ENGINE AIR INLET SCREENS RETRACTED.
♦OPTIMUM cruise altitude normal RATED power.
Figure A-15 (Sheet 1 of 3).
A-53
Appendix I
T.O. 1F-89H-1
MODEL: F-89H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
BEST CimS PERFORMANCE (Tim)
MILITARY POWER
PYLON TANK CONFIGURATION
ENGINE®: (2) J3545
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB/US GAL
GROSS WEIGHT - 1000 LB
REMARKS:
1. FUEL CONSUMPTION INCREASED 5 PERCENT FOR SERVICE VARIATION.
2. SUBTRACT 906 POUNDS FROM AVAILABLE FUEL TO AUOW FOR WARMUP.
TAXI, AND TAKEOFF; ENTER CHART AT TAKEOFF GROSS WEIGHT LESS
906 POUNDS,
3. ENGINE AIR INLET SCREENS RETRACTED.
♦OPTIMUM CRUISE ALTITUDE - NORMAL RATED POWER,
HB5
Figure A-15 fSJieef 2 of 21
A-54
T.O. 1F-89H-1
Appendix i
BEST cum PEREORMAHCE (me)
MODEL: F-89H NORMAL ROWER ENGINE(S): (2) J35-35
DATA BASIS' FLIGHT TEST PYLON TANK CONFIGURATION FUEL GRADE: JP-4
DATE: 22 OCTOBER 1957 FUEL DENSITY: 6.5 LB/US GAL
GROSS WEIGHT - 1000 LB
REMARKS:
]. FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION,
2, SUBTRACT 906 POUNDS FROM AVAILABLE FUEL TO ALLOW FOR WARMUP.
TAXT AND TAKEOFF, ENTER CHART AT TAKEOFF GROSS WEIGHT LESS
906 POUNDS,
3. ENGINE AIR JNlET SCREENS RETRACTED.
* OPTIMUM CRUISE ALTITUDE - NORMAL HATED POWER,
H336
Figure A-15 (Sheet 3 of 3J,
A-5 5
Appendix ]
TO. 1F-89H-T
MODEL: F-89H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER T957
BEST CUMB SPEED
MAXIMUM POWER
PYLON TANK CONFIGURATION
ENGINE(S): (2) J3M5
FUEL GRADE: JP-4
FUEL DENSITY:6.5 LB/US GAL
l
a
<
0 7 4 6 a TO 12
160 240 320 400 400
RATE OF CLIMB - 1000 FT. MIN
REMARKS:
BEST CLIMB SPEED - KNOTS CAS
1 CLIMB AT CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE
2 . ENGINE AJR tNLET SCREENS RETRACTED
H337
Figure A-Id (Sheet I of 3).
A-56
T.O. 1F-89H-1
Appendix 1
MODEL F-89H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
bcst cum speto
MILITARY POWER
PYLON TANK CONFIGURATION
EMGINEfS): (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6,5 LB/US GAL
REMARKS;
BEST CLIMB SPEED - KNOTS CAS
1, CLIMB AT CA$ SHOWN REGARDLESS OF AMBIENT TEMPERATUftE.
2 . ENGINE AIR INLET SCREENS RETRACTED.
Figure A-16 (Sheet 2 of 3}.
A-57
Appendix 1
T.O. 1F-89H-1
MODEL F-89H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
best cum spcbb
NORMAL POWER
PYLON TANK CONFIGURATION
ENGINES): (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB/US GAl
\
UJ
Q
RATE OF CLIMB - 1OO0 FT/MIN
130 *200 220 240 260
BEST CLIMB SPEED - KNOTS CAS
2B0 300
REMARKS;
1. CLIMB AT CAS SHOWN REGARDLESS OF AM&ENT TEMPERATURE.
2. ENGINE AIR INLET SCREENS RETRACTED
K339
Figure A-16 (Sheet 3 of 3).
A-58
T.O. 1F-89H-1
Appendix 3
MODELS F-89H
DATA BASIS: FUGHT TEST
DATE: 22 OCTOBER 1957
BEST cum PERFORMANCE (RANGE)
MAXIMUM POWER
BASIC CONFIGURATION PLUS PYLONS
ONE ENGINE OPERATING
ENGlNEfS): (2) J35-35
FUEt GRADE: JP-4
FUEL DENSITY: 6*5 LB/US GAL
I* FUEL CONSUMPTION INCREASED 5 PERCENT FOR SERVICE VARIATION.
3. SUBTRACT 906 POUNDS FROM AVAILABLE FUEL TO ALLOW FOR WARMUP.
TAXI. AND TAKEOFF; ENTER CHART AT TAKEOFF GROSS WEIGHT LESS
906 POUNDS.
3. ENGINE AIR INLET SCREENS RETRACTED.
* OPTIMUM CRUISE ALTITUDE - NORMAL RATED POWER
Figure A-17 (Sheet I of 2 ).
A-59
Appendix I
T.O. IF-89H-1
MODEL: F-89H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
BIST Cim MBFOBM/Utce (MMCt)
MILITARY POWER
BASIC CONFIGURATION PLUS PYLONS
ONE ENGINE OPERATING
ENGINE(S); (2) 335-35
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB/US GAL
GROSS WEIGHT - KXX) LB
REMARKS;
1. FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION.
2. SUBTRACT 906 POUNDS FROM AVAILABLE FUEL TO ALLOW FOR WARMUP*
TAXI, AND TAKEOFF; ENTER CHART AT TAKEOFF GROSS WEIGHT LESS
906 POUNDS.
3. ENGINE AIR INLET SCREENS RETRACTED,
* OPTIMUM CRUISE ALTITUDE - normal RATED POWER.
H34I
Figure A-17 (Sheet 2 of 2).
A-60
T.O. TF-89H-1
Appendix 1
MODEL: F-89H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
BEST CUM8 PERFORMANCE (Tim)
MAXIMUM POWER
BASIC CONFIGURATION PLUS PYLONS
ONE ENGINE OPERATING
ENGINE(S): (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB/US GAL
z
LU
5
GROSS WEiGHT —1000 IB
REMARKS
1. FUEL CONSUMPTION INCREASED 5 PERCENT FOR SERVICE VARIATION,
7 . SUBTRACT 006 POUNDS FROM AVAILABLE FUEL TO ALLOW FQft WARMUP,
TAXI, AND TAKEOFF; ENTER CHART AT TAKEOFF GROSS WEIGHT LESS
906 POUNDS.
3. ENGINE AIR INLET SCREENS RETRACTED,
* OPTIMUM CRUISE ALTITUDE NORMAL RATED POWER.
H342
Figure A-T8 (Sheet I of 2) m
A-61
Appendix I
T.O. 1F-89H-1
MODEL: F-89H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
BEST CUM8 PERFORMANCE (me)
MILITARY POWER
BASIC CONFIGURATION PLUS PYLONS
ONE ENGINE OPERATING
ENGINES).:® J35-35
FUEL GRADE: J P-4
FUEL DENSITY: 6,5 LB/US GAL
V)
21
Z
£
I
s
u
uh
5
4B 46 44 42 40 25 26 34 32 30
GROSS WEIGHT - TOGO LB
REMARKS:
1. FUEL CONSUMPTION INCREASED 5 PERCENT FOR SERVICE VARIATION.
2. SUBTRACT 906 POUNDS FROM AVAILABLE FUEL TO ALLOW FOR WARMUP,
TAXI, AND TAKEOFF; ENTER CHART AT TAKEOFF GROSS WEIGHT LESS
906 POUNDS.
3. ENGINE AIR INLET SCREENS RETRACTED
OPTIMUM CRUISE ALTITUDE - NORMAL RATED POWER.
H30
Figure A-18 (Sheet 2 of 21
A-62
T.O. 1F-89H-1
Appendix I
MODEL; F-89H
DATA BASIS; FLIGHT TEST
DATE: 72 OCTOBER 1957
b£$t cam spm
MAXIMUM POWER
BASIC CONFIGURATION PLUS PYLONS
ONE ENGINE OPERATING
ENGlNEfS)' (2) J35-35
FUEL GRADE; JP^4
FUEL DENSITY 6,5 LB/U5 GAL
0 CM 0.8 1.2 1.6
ft A T6 OF CLIMB - 1000 FT/MlN
2,0 2.4 2.8
180 200 220 240 260 280
BEST CLIMB SPEED - KNOTS CAS
REMARKS:
1- CLIMB AT CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE.
2. ENGINE AIR INLET SCREENS RETRACTED-
H34*
Figure A-19 (Sheet I of 2).
A-63
Appendix l
TO. 1F-B9H-1
MODEL; F-S9H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER T957
am cam spied
ENGINEfSb C2) J 35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB/US GAL
MILITARY POWER
BASIC CONFIGURATION PLUS PYLONS
ONE ENGINE OPERATING
BEST CUMB SPEED - KNOTS CAS
REMARKS:
I CLIMB AT CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE-
Z ENGINE AIR INLET SCREENS RETRACTEO-
Figure A-19 (Sheet 2 of 2).
A-64
TO. 1F-89H-T
Appendix I
NAUTICAL Atll£S PEN WOO POUNDS fUH
SEA LEVEL
MODEL: F-39H BAS | C CONFIGURATION PLUS PYLONS ENGJNE(S): (2) J 35-35
DATA BASIS: FLIGHT TEST FUEL GRADE: JP-4
DATE: 22 OCTOBER 1957 FUEL DENSITY: 6.5 LB/US GAL
160 200
240
TRUE AIRSPEED—KNOTS
280 320 340
400
480 S20
160 200
REMARKS:
240
280
320 360
CALIBRATED AIRSPEED—KNOTS
400
4B0 S20
1. FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION.
2 . ENGINE AIR INLET SCREENS RETRACTED.
3. MAINTAIN CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE.
H34&
Figure A-20 (Sheet 1 of 9),
A-65
Appendix !
T.O. 1F-89H-1
NAUTICAL MU£$ PER WOO POUNDS FUU
5000 FEET
BASIC CONFIGURATION PLUS PYLONS
ENGENE1S): (2) J3S-35
MODEL: F-B9H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
FUEL GRADE: JP-4
FUEL DENSITY: 6,5 IB/US GAL
MACH NUMBER
TRUE AIRSPEED-KNOTS
1 360
lift 200 24Q 2S0 320 360
160 CALIBRATED AJRSPEED-KNOTS
REMARKS;
1, FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION*
2, ENGINE AIR INLET SCREENS RETRACTED
3, MAINTAIN CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE*
K3*7
Figure A-20 (Sheet 2 of 9).
A-66
T.O. IF-89H-1
MODEL; F-89H
DATA BASIS; FLIGHT TEST
DATE; 22 OCTOBER 1957
NAUTICAL AUICS PBR 1000 POOHBS fOU
10,000 FEET
BASIC CONFIGURATION PLUS PYLONS
REMARKS:
L FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VACATION.
2 ENGINE AIR INLET SCREENS RETRACTED,
3. MAINTAIN CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE-
Figute A-20 (Sheet 3 of 9)
Appendix I
T.O. TF-89H-I
NAUTICAL Atll£S PER 1000 POUNDS FUEL
MODEL F-89H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
15,000 FEET
BASIC CONFIGURATION PLUS PYLONS
ENGINE S': (2) i35-35
FUEL GRADE: JP-4
FUEL DENSITY 6.5 LB/US GAL
200 240 280 320 360 400 440 400 520
1*0 200 240 280 320 360 400 440
CALIBRATED AIRSPEED-KNOTS
REMARKS:
1. FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION.
2 ENGINE AIR INLET SCREENS RETRACTED.
3 MAINTAIN CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE.
H349
figure A-20 (Sheet 4 of 9).
A-68
T.O. 1F-89H-1
Appendix I
nautical mas pc* mo pounds run
MODEL F-89H
20,000 FEET
BASIC CONFIGURATION PLUS PYLONS
ENGINE’SL (2) J35-35
DATA BASIS- FLIGHT TEST
DATE; 22 OCTOBER 1957
FUEl GRADE JP-4
FUEL DENSITY- 6.5 LB/US GAL
\. FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VACATION.
2. ENGINE A Eft INLET SCREENS RETRACTED.
3, MAINTAIN CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE,
H1S0
Figure A -20 (Sheet 5 of 9 J,
A-69
Appendix I
T.O. 1F-89H-I
NAUTICAL NUltS PtK 1000 POUNDS fUU
MODEL: F-89H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
25,000 FEET
BASIC CONFIGURATION PLUS PYLONS
ENGINE(S): {2} 135-35
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB /US GAL
220
rr
REMARKS:
260
300
340
1_!
TRUE AIRSPEED-KNOTS
380 420
II I III II I I
160 180 200 220 240 260 280 300
CALIBRATED AlRSPEED^KNOTS
2 . T ° AUOW F ° R SKV,CE VAR ' ATION -
3, MAINTAIN CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE
460
500
TTT
540
ITT
TTT
320 340 360 380
H35T
Figure A-20 fSheef 6 of 9).
A-70
T.O. 1F-89H-1
Appendix
NAUTICAL Mll€S PER 1000 POUNDS fUU
MODEL: F-89H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
30,000 FEET
BASIC CONFIGURATION PLUS PYLONS
ENGINE1S): (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB/US GAL
-maximum I upf
-ENDURANCE 4
GROSS WEI
§ 32,000 L
36,000 l
40,000 I
44,000 I
LONG RANGE
-CRUISE SPEED
-ZERO WIND
■As<fc?Pr t
i ift f fr
MACH NUMBER
TRUE AIRSPEED-KNOTS
440
240 260 2B0
CALIBRATED AIRSPEED—KNOTS
1. FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION
2. ENGINE AIR INLET SCREENS RETRACTED.
3. MAINTAIN CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE,
Figure A-20 (Sheet 7 of 9).
MACH NUMBER
TRUE AIRSPEED-KNOTS
380 420
CALIBRATED AIRSPEED—KNOTS
1. FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VACATION.
2. ENGINE AIR INLET SCREENS RETRACTED.
3. MAINTAIN CAS SHOWN REGARDLESS OP AMBIENT TEMPERATURE.
Figure A-2Q (SJieef 8 of 9),
T.O. 1F-89H-1
Appendix I
NAUTICAL mes P£R WOO POUNDS fOU
MODEL; F-S9H
40,000 FEET
BASIC CONFIGURATION PLUS PYLONS
ENGINE(S): (2) J35-35
DATA BASIS; FLIGHT TEST
DATE; 22 OCTOBER 1957
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB/US GAL
REMARKS:
280 320
TRUE AIRSPEED—KNOTS
360 400 440
400
III I
u
FT
140 160
100 200 220
CALIBRATED AIRSPEED-KNOTS
240 260
520
280
1. FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION,
2. ENGINE AIR INLET SCREENS RETRACTED,
3. MAINTAIN CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE.
Figure A-20 (Sheet 9 of 9),
A-73
Appendix I
T.O. 1F-89H-1
NAUTICAL PULES PER WOO POUNDS FUU
MODEL M9H
DATA BASIS: FLIGHT TEST
DATE; 22 OCTOBER T957
SEA LEVEL
PYLON TANK CONFIGURATION
ENGfNErSj: (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6*5 LB/US GAL
160
200
240
160
200
280
TRUE AIRSPEED—KNOTS
320 360
240
400
440
REMARKS:
> 320 360
CALIBRATED AlRSPEED—KNOTS
400
400
400
1 FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION.
2. ENGINE AIR INLET SCREENS RETRACTED,
3. MAINTAIN CAS SHOWN REGARDLESS Of AMBIENT TEMPERATURE,
H3S$
Figure A-21 (Sheet 1 of 81
A-74
T.O. 1F-89H-1
Appendix [
NAUTICAL M/L£S PEN WOO POUNDS FUEL
MODEL F-89H
5000 FEET
PYLON TANK CONFIGURATION
ENGINE'S;:(2) J35-35
DATA BASIS: FLIGHT TEST FUEL G8ADLJP-4
DATE: 22 OCTOBER 1957 FUEL DENSITY:6,5 LB/US GAL
MACH NUMBER
L FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION,
2 ENGINE AIR INLET SCREENS RETRACTED.
3, MAINTAIN CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE.
Figure 4-21 (S/ieef 2 of 8),
A-75
Appendix 3
T.O* 1F-89H-1
MACH NUMBER
1. FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION,
2. ENGINE AIR INLET SCREENS RETRACTED,
3. MAINTAIN CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE,
K3S7
Figure A-21 (Sheet 3 of 8),
A-76
T.O. 1F-89H-1
Appendix I
NAUTICAL MILES PEN 1000 POUNDS fUEL
15,000 FEET
PYLON TANK CONFIGURATION ENGINES): (2) J35-35
FUEL GRADE: JP~4
FUEL DENSITY LB/ US GAL
MODEL: F-B9H
DATA BASIS FLIGHT TEST
DATE: 22 OCTOBER 1957
REMARKS:
CALIBRATED AIR SPEED-KNOTS
1 FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION
2 ENGINE AIR INLET SCREENS RETRACTED
3, MAINTAIN CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE-
HIKl
F/gure A-21 (Sheet 4 of 8)>
A-77
Appendix I
TO. 1F-89H-1
nautical Maes pen mo pounds fuh
MODEL: F-S9H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
20,000 FEET
PYLON TANK CONFIGURATION
ENGINEfSh (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY 6.5 LB/US GAL
TRUE AIRSPEED—KNOTS
200 240 2B0 320 360 400 440 4S0 S20
160 200 240 ISO 320 360
CA1IBRATED AIR SPEED— KNOTS
REMARKS:
1. FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION.
2 ENGINE AIR INLET SCREENS RETRACTED,
3. MAINTAIN CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE.
Has?
Figure A-21 (Sheet 5 of 8^,
A-78
TO. 1F-89H-1
Appendix I
NAUTICAL MIUS PEN WOO POUNDS fUU
MODEL, F-B9H
DATA BASIS: FLIGHT TEST
DATEr 22 OCTOBER 1957
25,000 FEET
PYLON TANK CONFIGURATION
ENGINE 5 ; (2) J35-35
FUEL GRADE JP-4
FUEL DENSITY 6.5 LB/US GAL
z
3
o
&
s
«
Cl
s
<
K =L
<
2
MACH NUMBER
TRUE AIRSPEED-KNOTS
CALIBRATED AIRSPEED-KNOTS
REMARKS;
1. FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION,
7. ENGINE AIR INLET SCREENS RETRACTED,
3. MAINTAIN CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE.
Figure A-21 (Sheet 6 of 8).
A-79
Appendix I
T.O. 1F-89H-1
nautical mutes peg mo pounds ruu
MODEL; F-89H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1557
30,000 FEET
PYLON TANK CONFIGURATION
ENGlNE:S)r(2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY 6,5 L8/US GAL
MACH NUMBER
1. FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION.
2. ENGINE MR INLET SCREENS RETRACTED.
3. MAINTAIN CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE.
Hist
Figure A-2I (Sbeef 7 of 8).
A-80
T.O* 1F-89H-1
Appendix E
NAUTICAL MUCS NS 1000 POUNDS fuu
MODEL: F-89H
DATA BASIS: FLIGHT TEST
DATE' 27 OCTOBER 1957
35,000 FEET
PYLON TANK CONFIGURATION
ENGINE S ( 2 ) J35-35
FUEL GRADE JP^4
FUEL DENSITY 6,5 LB/US GAL
REMARKS:
300
ISO
m<E AIR SPEED—KNOTS
330 420 460
220 240 260 2SO
CALIBRATED AIRSPEED—KNOTS
3- FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION
2. ENGINE AIR INLET SCREENS RETRACTED.
3. MAINTAIN CAS SHOWN REGARDLESS Of AMBIENT TEMPERATURE-
v
H343
Figure 4-21 fS/ieet 8 of 8).
A-81
Appendix 9
T.O. 1F-89H-1
NAUTICAL MU£S PBN 1000 POUNDS FUU
MODEL: F-89H
DATA BASIS: FLIGHT TEST
DATE 22 OCTOBER 1957
SEA LEVEL
BASIC CONFIGURATION PLUS PYLONS
ONE ENGINE OPERATING
ENGINE'S: (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY 6,5 LB/US GAL
MACH NUMBER
REMARKS:
160 200
TRUE AIRSPEED—KNOTS
240 280 320 360
160
200 240 280 320
CALIBRATED AIRSPEED—KNOTS
360
1. FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION
2. ENGINE AfR INLET SCREENS RETRACTED
3. MAINTAIN CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE.
Figure A-22 (Sheet I of 4).
A-S2
T.O. 1F-89H-1
Appendix I
NAUTICAL MILES PER 1000 POUNDS FUEL
MODEL F-89H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
5000 FEET
BASIC CONFIGURATION PLUS PYLONS
ONE ENGINE OPERATING
ENGINES): (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6,5 LB/US GAL
* .3 .4 js * j
MACH NUMBER
REMARKS.
TRUE AIRSPEED-KNOTS
200 240 2SO 320 360
T60 200 240 280 320
CALIBRATED AIRS PEED—KNOTS
I FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION
2. ENGINE AIR INLET SCREENS RETRACTED,
3. MAINTAIN CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE.
K36J
Figure A-22 {Sheet 2 of 4).
A-83
Appendix I
T.O. 1F-89H-1
NAUTICAL Mtl£$ P€N W0& POUNDS FUH
10,000 FEET
MODEL F-89H ENGINE®: (2) J 35-35
BASIC CONFIGURATION PLUS PYLONS
DATA BASIS: FLIGHT TEST FUEL GRADE: JP-4
DATE: 22 OCTOBER 1957 ONE ENGINE OPERATING FUEL DENSITY 6.5 LB/US GAL
180
160
TRUE AIRSPEED—KNOTS
220 260 300 340
180 200 220 240 260 280 300
CALIBRATED AIRSPEED-KNOTS
REMARKS:
L FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION.
2. ENGfNE AiR INLET SCREENS RETRACTED,
3, MAINTAIN CA$ SHOWN REGARDLESS OF AMBIENT TEMPERATURE,
H3*5
Figure A-22 (Sheet 3 of 4).
A-84
T.O. 1F-89H-1
Appendix I
NAUTICAL MU£S PEN WOO POUNDS FUU
MODEL: F-89H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
15,000 FEET
BASIC CONFIGURATION PLUS PYLONS
ONE ENGINE OPERATING
ENGINEfSj: (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY 6-5 LB/US GAL
REMARKS:
true airspeed-kisigts
160 200 240 280 320
130 ISO 170 190 210 230 250 270
CALIBRATED AIRSPEED—KNOTS
1 FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION.
2. ENGINE AIR INLET SCREENS RETRACTED.
3. MAINTAIN CAS SHOWN REGARDLESS OF AMBIENT TEMPERATURE.
K*64
Figure A-22 (Sheet 4 of 4),
A-85
Appendix I
T,0. 1F-89H-1
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A-86
T.O. 1F-89H-1
Appendix I
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Appendix \
T.O* 1F-89H-1
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A-88
T.O. 1F-89H-1
Appendix 1
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Appendix 1
T.O. 1F-89H-1
Figure A-24 (Sheet 2 of 3J.
A-90
UNI OF BEST RANGE FOR CONST AN?. ALTITUDE FLIGHT 37,000
T.0, 1F-S9H-T
Appendix 1
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Figure A-24 (Sheet 3 of 3).
A-9T
OPTIMUM RETURN PgOf/lt
DATA BASIS; FlIGHT TEST TAKEOFF GROSS WEIGHT
DATE; 22 OCTOBER 1957 43,175 POUNDS MODEL: F-89H
Appendix E
T.O, 1F-89H-1
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MAXIMUM ENDURANCE
DATA BASIS: FLIGHT TEST TAKEOFF GROSS WEIGHT
DATE: 22 OCTOBER 1957 43 175 POUNDS MODEL: F-89H
4J,I« ENGINE(S): (2) J35-35
_ FUEL GRADE: J P-4
| altitude | CONFIGURATION: BASIC PLUS PYLONS FUEL DENSITY: 6.5 LB/US GAL
T.O. 1F-89H-1
Appendix E
Figure A-26 (Sheet I of 3).
A-95
SEA LEVEL 19b 195 ,30 «0Q
mxmm endurance
DATA BASIS: FLIGHT TEST TAKEOFF GROSS WEIGHT
DATE: 22 OCTOBER 1957 47355 POUNDS MODEL: F-89H
ENGINE(S): ( 2 ) J35-35
_ -_ FUEL GRADE: JP-^
Appendix i
1 . 0 , 1F-89H-1
Figure A-26 (Sheet 2 of 3J.
A-96
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A-97
OPTIMUM MM/MOM CNDURANCt PftOftU
DATA BASIS'FLIGHT TEST TAKEOFF GROSS WEIGHT B „ u
____ MODEL: F-89H
DATE: 22 OCTOBER 1957 43,175 POUNDS ENGtNEES): (2) J3S-3S
Appendix I
T.O. TF-89H-1
Figure A- 27 fSheet ! of 3).
A-98
Appendix I
Figure A-27 (Sheet 2 of 3)
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DATE: 22 OCTOBER 1957 «,17S POUNDS MODEL: F-89H
Figure A~27 (Sheet 3 of 3).
A-100
T.O. 1F-89H-1
Appendix t
MODEL F-89H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
DESCENTS
IDLE POWER
WITH OR WITHOUT PYLON TANKS
ENGINEI5I: (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB/US GAL
1. FQR MAXIMUM RANGE DESCENT, MAINTAIN 200 KNOTS INDICATED AJRSPEED (IAS)
2. FOR RECOMMENDED DESCENT, MAINTAIN 0J MACH NUMBER
{SPEED BRAKES OPEN OR CLOSED l
3. FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION,
4. ENGINE AIR INLET SCREENS RETRACTED.
__44.000 IS
—-32,000 LB
Figure A-28 (Sheet 1 of 2).
A-101
PRESSURE ALTITUDE - 1COQ FEET PRESSURE ALTITUDE - 1000 FEiT
Appendix 1
T.O. 1F-89H-1
MODEt: F-89H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
DESCENTS
IDLE POWER
WITH OR WITHOUT PYLON TANKS
ONE ENGINE OPERATING
ENGiNEfSJ: {2} J 35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB/US GAL
1. FOR MAXIMUM RANGE DESCENT, MAINTAIN 200 KNOTS INDICATED AIRSPEED ilAS).
2. FOR RECOMMENDED DESCENT* MAINTAIN 0.7 MACH NUMBER {SPEED BRAKES OPEN OR CLOSED).
3. FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION. -44 000 Lfl
4. ENGINE AIR INLET SCREENS RETRACTED, —-- 32 000 LB
5. SINGLE^ENCtNE DESCENTS NOT RECOMMENDED BECAUSE OF THE POSSIBILITY OF
"DUCT RUMBLE'" ON THE WINDMILLING ENGINE.
Figure A-2B (Sheet 2 of 2K
A-102
PRESSURE ALTITUDE —■ 1000 FEET PRESSURE ALTITUDE - 1000 FEET
LANDING D/STANCe
MODEL: F-89H WITH OR WITHOUT PYLON TANKS * ENGINES): (2) J35-35
DATA BASIS: FLIGHT TEST FUEL GRADE: JP-<I
DATE: 22 OCTOBER 1957 FUEL DENSITY: 6.5 LB/ US GAL
TO, 1F-S9H-1
Appendix l
Figure A-29 (Sheet l of 4).
A-103
LANDING D/STANCt TO CltAR SORT-OBSTAClt
MODEL: F-89H WITH OR WITHOUT PYLON TANKS* ENGINE(S); (2) J35-35
DATA BASIS: FLIGHT TEST FUEL GRADE: JP-4
DATE: n OCTOBER 1M7 FUEL DENSITY: 6.5 LB/US GAL
Appendix 1
T.Q. 1F-89H-1
Id 0001-313VXS8O Xi OS avlID 01 3DNV1S1Q IVIOX
Figure A-29 (Sheet 2 of 4),
A-104
REMARKSr 1 USE SPEED BRAKES AS NECESSARY TO MAfNTAlN 3 CHART D[STANCES AND AIRSPEEDS ARE BASED ON
APPROACH AIRSPEED AND FULLY OPEN SPEED BRAKES NORMAL operating PROCEDURE AND USE OF
AFTER TOUCHDOWN. DRY HARD SURFACE RUNWAY
7. USE 50 DEGREE PUPS, *, ENOfNE AIR INLET SCREENS EXTENDED.
4 WITH EMPTY FYION TANKS ONLY
T\0. 1F-89H-1
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A-105
LANDING DISTANCE TO CLEAN SOFT. OBSTACLE
MODEL F-89H ONE ENGINE OPERATING ENGINE(S): (2) J35-3S
DATA BASIS: FLIGHT TEST WITH OR WITHOUT PYLON TANKS * FUEL GRADE: JP-4
DATE: 22 OCTOBER 1957 FUEL DENSITY- 6,5 LB/US GAL
Appendix I
T.O. 1F-89H-I
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A-106
Appendix I
T.O, 1F-89H-1
MODEL F-89H
DATA BASIS: FLIGHT TEST
DATE; 22 OCTOBER 1957
COMBAT AUOWANCl CHART
MAXIMUM POWER
BASIC CONFIGURATION PLUS PYLONS
ENGINES): (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB/US GAL
REMARKS:
1. FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VARIATION.
2. ENGINE AIR INLET SCREENS RETRACTED.
3 EXHAUST TEMPERATURE LIMIT: 750*0
H3BB
Figure A-31 fSheef J of 3h
A-108
T.O, IF-89H-I
Appendix
MODEL; F-S9H
DATA BASIS: FLIGHT TEST
DATE: 22 OCTOBER 1957
COMBAT AUOWAHCl CHART
MILITARY POWER
BASIC CONFIGURATION PLUS PYLONS
ENGINEfS): (2) J35-35
FUEL GRADE: JP^J
FUEL DENSITY; 6,5 LB/US GAL
TIME - MINUTES
REMARKS:
T. FUEL CONSUMPTION INCREASED 5 PERCENT TO ALLOW FOR SERVICE VACATION.
2. ENGINE AJR INLET SCREENS RETRACTED,
3. EXHAUST TEMPERATURE LIMIT: 750°C,
Figure A-31 (Sheet 2 of 3).
A-109
Appendix I
T.O. 1F-89H-1
COMBAT ALLOWANCE CHANT
MODEL F-B9H
DATA BASIS: FLIGHT TEST
DATE: 72 OCTOBER 1957
NORMAL POWER
BASIC CONFIGURATION PLUS PYLONS
ENGINE(S): (2) J35-35
FUEL GRADE: JP-4
FUEL DENSITY: 6.5 LB/US GAL
TIME - MINUTES
1. FUEL CONSUMPTION INCREASED 5 PERCENT TO A HOW FOR SERVICE VARIATION.
2. ENGINE AIR INLET SCREENS RETRACTED.
3. EXHAUST TEMPERATURE UMIT; 680* t
Figure A-3I (Sheet 3 of 3J,
T.O. TF-S9H-1
Index
A-2—Autopilot
/
>
>
HF-11A
A
Page
A-2 Flight Compute* ...... .
indicator ..., .,,..**.«**.*,
operation ... *.. ...........
flying compass course at constant altitude
starting and ground check ..
selector switch .. .
Accelerated Stalls ....................
Acceleration ...
burst .... .*..
limitations
Accelerometer ................ * * * *.,,
Acrobatics , J .. # ,* * * .......,
A-C Voltmeter and Selector Switch .........
ADF Filter Switch . *.,„
Aft Cert ter-of-Gravity Fuel Movement .
Afterburaer System
control switches .. . .. „ ,,
demand switches , . .
operation ....
starring at high altitude....
warning lights ...... .. .
After Ejecriort ......
After Landing ............. .
After Takeoff—Climb.. ..........
hot weather procedures ................
touch-and-go landings ....
Agent Discharge Switch ..
Ailerons ....
and elevator trim switch
and rudder movement . .,.
Airbrake
- 4-17
- 4-18
.. 4-18
... 4-19
-.. 4-18
. 4-17
....... 6-2
. 5 - 1,545
. 7-1
- 5-10
>.,._ 1-50
......... 5-10
... . 1 - 32 , 1-33
. 4-8
.. 3-24
........ 1-9
- 1-16
. 1-16
. 7-3
.. 7-3
-- 1-16
....... 3-17
-.... 2-21
- 2 - 13 , 8-2
- 9-20
. 2-19
.. 1-52
____ 6-4
. 1-37
-.. -- 5-10
emergency system ...
emergency' valve handle ...
Air-Gondltfonmg System .................
Air-Conditioning System* Cabin ..........
Airplane
dimensions ........... ...............
gross weights....
Airspeed
indicators ....
limitations ....
autopilot .....
landing gear . ......
landing—taxi light ...
pylon .... .......... *
pylon tank jettison ..
tire ..............................
wing flap........
Alternate Fuel Limitations..
Alternator Control Ponds ................
Alternator System ..... .................
a-c voltmeter and selector switch.
circuit breaker switch and indicator light
exciter switch ...
external power switch .................
failure ......
voltage rheostat ......................
Altimeter .............................
Altitude lost During Dive Recovery ....... .
Altitude Start and Starter-Test Switches . ..
Anti "G" Suit Equipment ...
Anti-Icing Control Panels ... . . .
Anti-Icing Switch . ......
Anti-Icing System Operation .............
descent *
in flight ^.
landing ...
takeoff ..
Anti-Icing Systems ....
fuel filter, low pressure..
radome ..
thermal and electrical ^..
windshield .......
Anti-Icing Warning Light ..
Approaches
autopilot-controlled—ILS ..
GCA........
ILS _ ......... ....
instrument ...
and letdowns on single engine...
radio ...
Approaching the Storm .....
Approach to Pattern .......
Armament ....
Armrests ...
ejection seat right . *..................
Asymmetrical Loading, Flight with .......
Asymmetrical Tip Fuel Condition VS Airspeed
Attitude Indicator ....
AttUotfe Indicator
Augmenter System, Sideslip Stability.
Automatic Approach Equipment .........
approach switch...
localizer
operation
Automatic Pilot Control Panel ............
Automatic Release, Safety Belt ..
Autopilot .......... 4 ..
(Boldface type denotes illustration}
Page
1-48
1-48
4-3
4-1
. 1-1
...... l-l
. 1-50
. 5-7
. 5-9
..... 59
..... 59
- 5 9
_ 5-10
..... 5-9
.. 5-9
. 5-7
_ 1-30
__ 1-32
...... 1-33
...... 1-33
. 1-32
....... 1-32
...... 3-26
. 1-33
- 1-50
6-10 - 6-15
...... 1-8
.. 4-32
___ . 4-6
...... 4*5
. 445
...... 4-6
. 4-6
. 4-6
...... 4-6
.. 4-5
- 4-7
...... 4-7
_ 4-5
.. 4-7
.. 4-6
...... 9-10
. 9-7
. 9-10
...... 9-7
...... 9-13
.* 9-7
...... 9-16
.. 9-19
... 1 - 1 , 4-29
.. 1-60
.. 1-56
...... 6-16
.3-18
- 1-50
.1-50
_ 1-40
. 4-29
. 4-29
__ _ 4-29
___ 4-29
...... 4-26
.. 1-60
.. 4-26
1
Index
Autopilot—Communication
T.O. 1F-S9H-1
Page
altitude switch 4-27
autotrim switch and indicator 4-27
check ......« . ( .2-10
disengaging procedure. ...... .4-28
emergency disconnect switch ............. . ,, , 4^27
engaging procedure in turns or uncoordinated flight. - 4-28
normal ........ ,, 4-28
engaging switch ...*.. 4-26
ground tests ...... . 4^28
heading trim indicator and knob ... 4-27
ILS—autopilot-controlled approach .. 9-10
limitations ...... . t 5 _c)
maneuvering flight. 4-28
operation
emergency . ..* 4 ,. .. 4-29
normal .. 4-28
pitch control knob .... 4-27
power switch .. 4-26
roll trim knob ..... 4-27
straight and level flight .............. .......... . 4-28
trimming procedure ............................ 4-28
turn knob....... t . 4-27
Auxiliary Equipment , .. 14>6
Axial-Flow Turbojet Engine . 1-6
1
Battery Switch ....«. 1-25
Before Ejection . 3-16
Before Entering Cockpit/Airplane ..... 2-2
cold weather procedure .... ..... 9-16
desert procedure - . ........................ ..... 9-21
hot weather procedure ........ 9-20
radar observer’s duties .., ..... 8-1
Before Exterior Inspection ..... 2-1
Before Landing .... ..-.. 2-15, 8-2
after touch-and-go ....... 2-21
Before Leaving Airplane ............... .. 2-22
cold weather procedure .. 9-20
desert procedure ..... .......... 9-2 1
hot weather procedure _______._ ..... 9-20
radar observer’s duties ................ .. . 8-2
Before Starting Engines...... 2-7, 9-17
Before Takeoff ...... 2-10
cold weather procedure ..... 9-18
desert procedure_____9-21
preflight airplane check . ..... 2-10
preflight engine check_____ 2-11
radar observer's duties ........___....... 8-2
Blind Flying Curtain Assembly.... 4-32
Booster Pumps ...... * *___ ......... 1-17
failure
main tank . ..... .............. 3-22
wingtank. .... 3-22
Both Engines Inoperative ... 3-9
Brake Hydraulic and! Air Systems .................... 1 -49
Brake System_ ............... 1-48
brake pedals ..._ 1-48
emergency airbrake .........._ ...... 1-48
emergency airbrake valve handle ..._ ........ 1 -48
emergency operation. 3-30
operation . ........ 7-3
parking brake lever . ...*.*__ .... 1-48
Buffet-1 "G’ T Flight. 6-6
Burst Acceleration . . 7-1
C
Cabin Air-Conditioning Control Panels __ .......... 4-2
Cabin Air-Conditioning System.. ........... 4-1
air switch...... 4-1
air temperature switch.... 4-2
differential pressure switch 4-1
emergency operation _ 4-2
normal operation ..... 4-2
pressure regulator _______4-1
temperature rheostat ....... 4-2
Page
Cabin Pressure Schedules .. 4_4
Canopy .,... I -54
defogging system .. 4.2
knobs * *.. ....... . 4-2
operation .. 4-4
ejection seat right armrest ..... j.gg
ejector pressure gage .. 1-54
external canopy
handgrips. 1-55
emergency release handle ... . 1.35
switches ..^ . l- 5 >
jettison system ....... 1 - 54 , 7*6
limitations .. g.p)
lock levers and indicator light ..... 1-55
loss of. 3.30
pilot's
handgrips . ........ 1^55
jettison T ’T" handle .... 1-56
switch ..... 1-54
radar observer's
handgrips .... 1-55
switch ... p , , 1-55
Canopy Controls .......... * * . . , ..... ...... . 1 -S 3
Catapult Firing Trigger ... 1-60
CenteMf-Gravity Limitations ....................... 5-15
Check
autopilot ........ 2-10
flight computer starting and ground.. . 4-18
hydraulic system .. 2-9
interior .... 2-5
radar observer’s ..... 8-2
oxygen system preflight ..... 4-24
preflight airplane ... 2-10
preflight engine ..*... 2-11
VHF navigation set ground ... 4-16
voltage ....... 2-9
Checklists .... 4-32
condensed........ 2-25, 3 - 31 , 8-5
Circuit Breaker Panels , *. 1-31
Circuit Breakers, 28-Volt D-C ...... 1-29
Climb ........... 2-14
instrument . 9-4
maximum distance ........ 2-15
maximum rate of ..— .... 2-14
minimum distance ..... . 2-15
minimum fuel ...... . 2-15
Climb with Might Computer ., 9-3
Cockpit
before entering... 2-2
lighting
C-4 cockpit lights......4-22
pilot's .. 4 - 2 !
radar observer's ..... ....... 4-21
rear .. 8-1
Cold Weather Procedures ....9-16
approach to pattern ......-........ ____ 9-19
before entering cockpit .. 9-16
before leaving airplane .. 9-20
before starting engines ...... 9-17
before takeoff ............ 9-18
during flight *...... .... 9-19
ground tests . 9-18
landing .. 9-10
starting engines .......... ...... 9-17
takeoff ...... .... 9-19
taxiing instructions.. ...... 9-18
Command Radio ...... 4-1 3
controls ,.... 4-13
operation .... ---- 4-13
Cemmond Radio Control Panel ...................... 4-13
Communications and Associated Electronic Equipment... . 4-9
Communication and Associated Electronic Equipment ... 4-8
A -2 flight computer..... 4-17
command radio . .. 4-13
glide-slope receiver .. 4-17
IFF .... 4-19
interphone .. 4-8
(Boldface type denotes illustration]
2
Changed 13 February 1959
TO. 1F-89H-I
Index
Com muni cat ion'—Eng i nes
marker beacon receiving , . *.* ■ * *
radio compass.
YHF navigation .
Compass, Radio .
Compressor Stall .
Condensed Checklist...
Continued Flight Impossible .
Control Stick .
Control Stick Grip ..
Cooling and Air Induction System, Engine
Course Indicator ♦...
C«ursQ Indicator ....
Crew Requirements, Minimum ..
Crossfeed Operation *,.,.*******
Crossfeed Switch .
Crosswind
landing ...* * * * *
takeoff ...
Cruise.
Cruising and High Speed * * - *..
Cruising Flight, Instrument.-
Curtain Assembly, Blind Flyipg.-
D
Damaged Tanks ..* * * * *
main....
tip or pylon .
wing.
Danger Areas....
Data Cases, Map and ..
D-C Control and Indicator Panel, 28-Volt
D-C Generator Malfunction Chart, 28-Volt
D-C System, 28-Volt.
battery switch.
circuit breakers.
generator switches .
generator warning lights.
loadmeters ...
voltage regulator rheostats .
voltmeter and voltmeter selector switch .
Defogging System, Canopy..
De-Icing System, Fuel Filter, low pressure
Descent ..
instrument *.* *. *.
Descent with Flight Computer . *.
Desert Procedures .
before entering cockpit.-
before leaving airplane *****.**.**»,<
before takeoff..
takeoff ..
Detachable Parachute Lanyard ****•**<-»
Dimensions ..
Dimensions, Airplane .
Dimming Switch, Warning lights.
Ditching ...
Diving , .*..
high Mach dive .
During Flight
cold weather procedures.
radar observer’s duties.
Duties
pilot's .*..
radar observer’s .
E
E-9 Fire Control System *, * *,
Ejection .
after ..
before.
failure of canopy to jettison
failure of seat to eject .....
procedure .
Ejection Procedure * - *.*
Ejection Seats .
armrests = -.
right.
Page
.4-17
. 4-14
.4-14
. 4-14
. 7-1
2-25,3-31,8-5
. 3-5
.... 1-37
.1-37
. 1-6
.4-15
.4-15
. 5-1
. 7-3
. 1-19
. 2-18
....2-13
.2-15
. 6-5
. 9-4
. 4-32
*... 3-23
*,*, 3*23
.... 3-24
_3-24
_ 2-8
_4-32
_1*28
_3*25
*,** 1-25
- 1-25
**** 1-29
_ 1-28
...* 1-29
- 1-29
- 1-25
_ 1-29
- 4-2
**1-17,4-7
_ 2-15
*..* 9-7
_ 9-6
- 9*21
_9*21
.... 9*21
_9-21
_9-21
,... 1-61
_ 1-2
* *. * M
-4-21
_ 3-20
*.. * 6-7
* * * - 6-8
**** 9*19
_ 8-2
8*1
8*1
.4-29
.3*13
- .3*17
.3-16
. 3*17
.3*16
.3*16
3-14-—*3-15
- . 1-56
. 1-60
.1-56
catapult firing trigger.
ground safety pins.
low altitude “one and zero" ejection system
safety belt automatic release * *.* ♦
seat adjustment lever.* * * *.
shoulder harness inertia reel lock lever-
Ejection Seats .*.
Ejector Pressure Gage, Canopy.
Electrical Fire, Fuselage, Wing, or.
Electrical Power Distribution * * *.* * * * *
Electrical Power Supply Systems.
alternator system ...*
a*c voltmeter and selector switch . * *-
circuit breaker switch and indicator light
exciter switch ....*
external power switch..
voltage rheostat -_.
electrically operated equipment.
external power system.
inverter systems .
a-c voltmeter selector switch .
single-phase inverter switch .
single-phase inverter warning light *. * * *
three-phase inverter switch —.
three-phase inverter warning light.
d-c system, 28-volt.
battery switch.
circuit breakers.
generator switches .
generator warning lights.
loadmeters .
voltage regulator rheostats...
voltmeter and voltmeter selector switch *
Electrical Rudder Trim Knob * * * *...
Electrical System Emergency Operation.
alternator failure .
generator
failure....* * *
overvoltage .
instrument failure .
engine . ...... .*
flight .
inverter failure.
Electrical System Lead Distribution Table * * *.
Elevator .
feel system...
trim position indicator.
Emergencies, Landing ...
Emergency
airbrake system.
valve handle.
entrance ...
equipment ...
exit on ground.
landing gear system.
signal system..
Emergency Fuel Flow...
Emergency Operation, System and Equipment
autopilot .
brake ..
cabin air-conditioning ., *. *.* - *.
electrical .*. *...*♦*.
flight control .
fuel.
hydraulic .
IFF .
landing gear ..
oxygen .*.
sideslip stability augmenter.
speed brake.
wing flap .
Emergency Override Lever Operation .
Engint Fuel Control System ..
Engine Fuel Control System.
Engines.
burst acceleration..
compressor stall .
cooling and air induction system .........
(Boldface type denotes Illustration)
Page
. 1450
. 1-66
. 1-63
. 1-60
. 1-60
. 1-66
1-57—1-58
.. 1-54
.3*13
1-26—1-27
. 1-23
. 1-32
. 1-33
. 1*33
. 1*32
. 1*32
. 1*33
. 1-25
. 1-25
. 1*29
. 1*32
. 1-30
.1*32
..1-32
. 1*32
. 1*25
. 1-25
. 1-29
.1-28
. 1-29
. 1-29
. 1*25
.1*29
. 1*38
. 3 -24
.3*26
. 3-24
.3*24
. 3*26
. 3-26
. 3-26
. 3-26
.1-24
. 6*2
. 1*37
. 1-40
.3*17
. 1-48
. 1-48
.3-20
. 1-52
. 3*20
. 1-44
.4-32
. 7-5
.4*29
.3-30
. 4-2
.3-24
.3-27
. 3-22
.3-27
.4-20
.3*28
..4-25
.3*28
.3*28
.3*28
. 3-4
. 1-7
. 1-2
.. 1-2—7*1
. 7-1
. 7-1
. 1*6
3
Index
Engines — Flight
T.O. 1F-89H-1
Page
Page
engined riven fuel pump failure warning light . 1*9
exhaust gas temperature gages. 1-9
exhaust gas temperature variation. 7*1
eyelid operation .* - 7-2
failure .....* 3-1
during flight. 3-6
during takeoff (after leaving ground). 3*4
continued flight impossible. 3-5
takeoff continued ..* - - * 3*4
during takeoff (before leaving ground). 3*3
engine-driven fuel pump.-.- 3-23
landing with both engines inoperative. 3-9
landing with oae engine inoperative.* * 3-8
left . 3*9
right. 3-8
maximum glide . 3-7
restarting engine in flight.*. 3-6
simulated forced landing ...*. 3-9
simulated single-engine flameout *. 3-9
single-engine
flight characteristics. ............... 3-1
go-around . ............. . 3-9
procedure.. 3-1
takeoff . 3*9
fire
during flight .. 3-13
during start ... 3-13
fire selector switches.1-19,1-52
fuel control system. 1*2
throttle friction lever. 1-6
throttles . 1-2
fuel flowmeter indicators. 1-9
ground operation .. 2-9
icing . 9-14
in above freezing air temperature ......... ..... 9*14
in below freezing air temperature ......... ..... 9-14
indication of. 9-14
instruments .-.. 3-26
limitations . 5-1
acceleration.5-1,5-6
alternate fuel. 5-7
exhaust gas temperature versus ambient temperature 5-7
Starting. 5-1
oil pressure gages. 1-9
overspeeding at altitude. 7-2
overtemperature versus engine life. 7-2
preflight check. 2-11
screens . 1-6
starting . 2-7
cold weather procedure.* .9-17
left . 2-8
right . 2*9
starting and ignition system. 1*6
altitude start and starter-test switches. 1-8
starter and ignition switches ... — .. 1-8
Starting power switch. 1-9
stopping. 2-22
tachometers . 1-9
Entering Cockpit/Airplane
before. 2-2
cold weather procedure.9-16
desert procedure. 9-21
hot weather procedure.9-20
radar observer's duties ........................ 8-1
Entrance . 2-4
Entrance .. 2-2
emergency . 3-20
Equipment
anti '"G" suit. 4-32
automatic approach . 4-29
auxiliary .. 1-66
electrically operated . .. 1-25
emergency . 1*52
lighting . 4-20
miscellaneous ................................. 4*31
navigation, radio and./_ 9-7
Exhaust Gas Temperature Gages .. 1-9
Exhaust Gas Temperature Variation . 7-1
Exhaust Gas Temperature VS Ambient Temperature..... 5-8
Exhaust Gas Temperature Versus Ambient Temperature. 5-7
Exit on Ground, Emergency.... ..... 3-20
Exterior Inspection .* • * - 2-^—2-3
Exterior Inspection. 2-1
before . 2*1
radar observer's duties. 8-1
Exterior lighting. 4-20
landing-taxi light and control switches. 4*20
position lights and control switches.4-20
External Canopy Handgrips .. .. ■ - ♦ 1-55
External Canopy Switches .. ..... 1-55
External Emergency Canopy Release Handle. 1-55
External Loads, Flight with. 6-16
External Power System. 1-25
External Stores Emergency Release Handle. 1*19
Extinguishing System, Fire. 1-52
Eyelid Operation . 7-2
F
F-89H Scorpion.....
Failure
alternator.
engine .
generator ..
instrument...
engine .
flight .
inverter .
of canopy to jettison .
of seat to eject.-
oil system ...
pumps
engine-driven fuel .
main tank booster.
wing tank booster .......................
Feel System, Elevator.
Filter De-Icing System, Fuel.
Fire .
engine
during flight.
during start.
fuselage, wing, or electrical.
Fire Control System, E*9.* ■
Fire Extinguishing System.
agent discharge switch.
engine selector switches.
fire and overheat warning lights and test switch
Fire Extinguishing System...
Firing
catapult trigger.
rocket/missile .
Flap System, Wing .
flap operation.
limitations .
Flat Tire, Landing with.
Flight
buffet—1 "G".
characteristics ....
single engine.
during
cold weather procedures.
radar observer's duties.
in icing conditions.
instrument cruising .
instruments .
inverted .
maneuvering...
planning .. —.
restrictions ....
straight and level ..... ...
with asymmetrical loading .
with external loads ..
Flight Computet, A-2.
indicator .... ♦....
missed approach with ILS.
operation .
iv
.... 3-26
... 3-1
... 3*24
... 3*26
3-26
... 3-26
... 3-26
... 3-17
... 3-16
... 3-20
... 3-23
... 3-22
... 3-22
... 1-37
... 1*20
... 3*13
... 3-13
... 3-13
... 3-13
... 4-29
... 1-52
... 1-52
1-19,1-52
... 1-52
... T-S1
...1-60
... 5*10
... 1-41
... 7-6
... 5-9
... 3-20
... 66
... 2-15
... 3-1
.. 9-19
.. 8-2
.. 9*14
.. 9-4
.. 3-26
.. 5-10
4-28,66
.. 2-1
.. 2-1
.. 4-28
.. 616
.. 6-i6
.. 4-17
,. 4*18
9-13
(Boldface type denotes Illustration]
4
T.O. 1F-89H-1
Index
Flight — Hydraulic
Page
flying compass course at constant altitude.. ♦ * 4-19
starting and ground check .. 4-18
Selector switch ... 4-17
Flight Computer Indicator .*. 4-18
Flight Computer Selector Switch.... 4*13
Flight Control Hydraulic System ... . 1*36
Flight Controls . *.6-2
ailerons .. * ..* > • • .. 6-4
elevator . ....... ...... 6-2
fH G ±h overshoot . ...... ....**.*.****** 6-4
rudder .. ......... ...., 6-4
speed brakes . .,«i ****««*••••>> ... 6-4
trim .........................., * *., • * * • * 6-5
high airspeed overtrim . *....... ...* - * 6-5
Flight Control System ...... 1-35
control stick ... ... t *, 1-37
elevator feel system ......, . 1-37
emergency operation ......... . 3-27
rudder pedals . 1-37
adjustment crank ..... 1-37
trim system ...... 1-37
aileron and elevator trim switch ................ 1-37
electrical rudder trim knob .... 1-38
elevator trim position indicator .. 1-40
rudder trim switch... 1-38
Flight Control Trim System .. ..1-33
Flying Compass Course at Constant Altitude ......... 4-19
Flying, Night . . 9-16
landing .... ....................... 9-16
takeoff . 9-16
Forced Landing. 3*19
Forced Landing .. . 3*12
Friction Lever, Throttle .. 1-6
Fuel Control Pone! . ..... 1 -22
Fuel Quantity Data ........ , 1-18
Fuel Quantity Gages 1-23
Fuel System....1-20—1-21
Fuel Supply System .... . . 1-17
alternate fuel limitations ...... 5-7
booster pumps . 1-17
crossfeed switch 1-19
emergency operation .. 3-22
engine-driven pump failure warning lights ........ 1-9
engine fire selector switches ..... 1-19
engine fuel control system .. 1*2
external stores emergency release handle ........... 1-19
filter de-icing system, fuel, low r pressure ......... 1-17, 4-7
fuel flowmeter indicators .. 1-9
fuel selector switches ...... .... 1-19
operation ....... 7-3
crossfeed .. 7-3
p>Ion tank jettison system . 1-18
button .. 1-19
quantity gage and selector switch . . 1-22
single-point system ............... 3-19, 4-29
system warning lights. . 1-22
throttle-actuated fuel shutoff switches .. 1-19
tip tank fuel dump system.. 1-18
dump button . 1-19
Fuel System Emergency Operation. 3-22
afc center-of-gravity fuel movement. 3-24
booster pump failure
main tank ........ . 3-22
wing tank . . . ....... 3-22
damaged tanks . 3.23
main. 3-23
tip or pylon....... 3-24
wing .... ..... 3-24
engine-driven fuel pump failure .. 3-23
following complete electrical failure .. 3-22
gravity feed ...... . 3-23
tip tank not feeding ... 3-23
Fuel Vent System Malfunction 3-21
fuel level control shutoff valve malfunction ........ 3-22
fuel overboarding during climb or dive ... 3*21
Fume Elimination, Smoke and .. 3-13
Fuselage, Wing, or Electrical Fire . ... 3-13
Q
Gages
canopy ejector pressure ......................... 1 -54
exhaust gas temperature 1-9
fuel quantity and selector switch.. 1 -22
hydraulic system pressure ........................ 1-35
oil pressure .. ...... 1-9
oxygen system pressure and flow indicator ... 4-24
GCA Approach ... .... . 9-11
GCA Approach ................................ 9-7
Gear Fails to Extend
gear fails to extend because of mechanical binding ... 3-30
on emergency procedure , . .....3-29
on normal procedure ... - . 3-29
General Arrangement ... ■ 1 -4—1 *5
Generator
failure .... ............................ . 3-24
overvoltage .. 3-24
switches .. 1-28
warning lights .. 1-29
Glide, Maximum .. 3-7
Glide-Slope Receiver . ...... 4-17
Go-Around.. 2-19
missed-approach procedure ....................... 9-13
single-engine . 3-9
Go-Around ...................................... 2-20
"G” Overshoot......* - 6-4
Gravity Feed .. ........ .. ... 3-23
Gross Weight, Airplane . ..... 1-1
Gross Weights .... 5-14
Ground Check, VHF Navigation Set .. 4-16
Ground Locks, Landing Gear ...1-44
Ground Operation, Engine ^. 2-9
Ground Safety Locks .. 1 -44
Ground Safety Pins, Ejection Seat .... ..... 1-66
Ground Tests ..... . 2-9,4-28
autopilot check .. 2-10
cold weather procedure ......................... 9-18
hydraulic system check .. 2-9
radar observer’s duties .. 8-2
voltage check .. 2-9
H
Handgrips
canopy
external ....*....... 1-55
pilot's ....... 1-55
radar observer's .. 1-55
Heading Trim Indicator and Knob ... 4-27
Heat System, Windshield . 4-7
Heavy Weight Landing ^. 2-19
High Airspeed
cruising and .. 6-5
overtrim .. 6-5
wing drop ....... 6-6
High Mach Dive . ......6-16—6-17
High Mach Dive ................. *, 6-5
Hot Weather Procedure ..... 9-20
after takeoff—climb ...*.. 9-20
before entering airplane 9-20
before leaving airplane. 9-20
landing . . . . ,.. 9-20
takeoff . 9-20
Hydraulic Power Supply System . ... 1-34
Hydraulic Power Supply System . 1-33
hydraulic system operation. 7-6
speed brake operation ........................ 7-6
wing flap operation 7-6
hydraulic system pressure gages .. 1-35
left hydraulic system . 1-33
right hydraulic system .. 1-35
system emergency operation. 3-27
systems check .. 2-9
(Boldface type denotes illustration)
5
Index
Ice—Level
T.O. 1F-89H-1
I
Page
Page
Ice and Rain. 9-13
engine icing. 9.14
in above freezing air temperature. 9-14
in below freezing air temperature. 944
indication of engine icing . 9-14
flight in icing conditions *. 9-14
surface icing. 9-14
IFF .4-19
controls. 4-19
emergency operation . .. 4*20
normal operation.4-19
Iff Control Pone I. 4-20
IFR
interceptions .,, .. 9-4
Ignition System, Starting and. 1-6
1LS
auto-pilot controlled approach ♦.,,.* * *.. 9-10
flight computer missed approach with.. 9-13
IL$ Approach With Flight Computer.9-12
IL5 Approaches.9-10
inbound to outer marker.9-10
outbound . 9*10
outer marker and inbound on approach ..9*10
procedure turn. 9-10
Indicators
airspeed . 1-50
attitude. 1-50
autotrim.4*27
elevator trim position. M0
flight computer. 4-18
fuel flowmeter. 1-9
heading trim. .. 4-27
landing gear . M6
oxygen system flow. 4-24
VHF navigation set ,.. 4*15
course.4-15
radio magnetic. 4-15
Inertia Reel Lock Lever, Shoulder Harness. 1-66
Inspection, Exterior.2*1, 8-1
Instrument
approaches. 9-7
climb . 9-4
cruising flight. 9-4
descent . 9-7
failure. 3-26
engine. 3-26
flight . 3-26
letdowns and approaches on single engine. 9-13
panel vibrators . 1-4$
takeoff . 9-4
Instrument Markings.5-2—*5-6
Instruments . 1-48
accelerometer . 1-50
airspeed indicators . 1-50
altimeter. 1-50
attitude indicator. 1-50
engine . 3*26
flight . 3-26
instrument panel vibrators . 1*48
machmeter. 1-50
Instrument Takeoff With Flight Computer. 9-2
Interior Check .2-5,8-2
front cockpit. 2-5
rear cockpit. 8-2
Interior Lighting .. 4-21
pilot's and radar observer’s C-4 cockpit lights.4-22
pilot's cockpit lighting.4*21
rheostats.4*21
radar observer's cockpit lighting. 4-21
rheostats.4*21
warning lights dimming switch. 4-21
Interphone Control Panel .. 4-13
Interphone System. 4-8
ADF filter switch. 4*8
control panel. 4-8
operation .. 4-13
pilot’s microphone switches 4-8
radar observer’s microphone buttons. 4*8
Inverted Flight .. 5-10
Inverter Control Panel. .... 1-29
Inverter Systems. 1-29
a*c voltmeter and selector switch .. 1-32
inverter failure. 3*26
single-phase inverter switch.. 1-29
single-phase inverter warning light. 1-32
three-phase inverter switch . 1-32
three-phase inverter warning light ................ 1-32
J
Jettison Systems
canopy.1*54,7-6
pylon tank .... 1-18
L
Landing.2-18,5*10
after . 2*21
before.2-15,8*2
cold weather procedure. 9-20
crosswind.2-18
heavy weight. 2-19
hot weather procedure. 9*20
minimum run . 2-19
night .9-16
normal . 2-18
wet or icy runway. 2-19
Landing Emergencies. 3-17
forced. 3-19
one dp rank containing fuel .. 3-17
running off runway . 3-19
with both engines inoperative. 3*9
with flaps and speed brakes retracted.3-18B
with flat tire. 3-20
with gear partially extended . 3-20
with lateral unbalance and critical aft CG. 3-17
with one engine inoperative. 3-8
left . 3-9
right. 3-8
Landing Gear Controls.1-45
Landing Gear Hydraulic System.1-42
Landing Gear System... 1-42
emergency operation ..3-28
gear fails to extend
gear fails to extend because of mechanical binding 3*30
on emergency procedure. 3-29
on normal procedure. 3-29
emergency override lever .. 1-46
emergency release handle. 1*46
emergency system. 1-44
ground locks. 1-4 4
lever.*. 1-44
limitation. 5-9
position indicators.... 1-46
warning horn and reset button .. 1-46
Landing Pattern ..3-16—5-17
Landing-Taxi Light
control switches. 4-20
limitation .. 5-9
Leaving Airplane, Before. 2-22
cold weather procedure.9*20
desert procedure. 9-21
hot weather procedure.9-20
radar observer’s duties. 8-2
Left Engine. 2-8
inoperative. 3-9
Left Hydraulic System. 1*33
Letdowns and Approaches on Single Engine, Instrument. 9*13
Level Flight Characteristics. 6-5
buffet—I "G” flight. 6-6
cruising and high speed. 6-5
high airspeed wing drop. 6-6
(Boldface type denotes illustration)
6
Changed 13 February 1959
Index
Level—Pilot's
T.0< 1F-89H-T
Page Pag?
low speed ............... 6*5
Lighting Control Pa nets. 4-20
Lighting Equipment ,.. ...... ... 4-20
exterior lighting , .. .. 4-20
landing-taxi tight and control switches ........... 4*20
position lights and control switches ............. 4-20
interior lighting .... . 4-21
pilot's and radar observer's C-4 cockpit lights ..... 4*22
pitot's cockpit lighting.4-21
rheostats ... 4-2 1
radar observer's cockpit lighting. 4-21
rheostats . ..... . 4-21
warning lights dimming switch. 4-21
Limitations
acceleration . 5-10
airspeed .. . 5-7
autopilot .. 5-9
landing gear ................................ 5-9
landing—taxi light... 5-9
pylon . 5-9
pylon tank jettison. 5-10
tire . 5-9
wing flap .. 5-9
canopy . 5-10
center*of-gravity . 5-15
engine. 5-1
acceleration. 5 -l 3 5-6
alternate fuel..... 5-7
starting . 5*1
weight . 5-15
Load Factors ..................................... 6*7
Loadmeters, 28-Volt D-C .. 1*29
Loss of Canopy ... 3-30
Low Pressure Fuel Filter De-Icing System ......... 1-17,4*7
Low Speed.. <5-5
M
Machmeter. 1-50
Mock Number Chart. 6*2
Main Differences Table .. 1*3
Main Tank
booster pump failure. 3-22
damaged. 3-23
Maneuvering Flight.4*28,6-6
load factors . 6-7
stick forces .. 6-6
Maneuvers, Prohibited. 5-10
Map and Data Cases.... .,, 4*32
Marker Beacon Receiving Set. 4-17
Maximum
distance climb. 2-15
glide. 3.7
rate of climb. 2*14
Maximum Glide ... h .. 3-7
Maximum Weights far Continued Flight After Engine
Failure an Takeoff . .. 3-5
Minimum
crew requirements . 5-1
distance climb. 2-15
fuel climb., 2-15
run landing.2-19
run takeoff. 2*13
Mirrors, Rear View .. 4-32
Miscellaneous Equipment.4-31
anti H *G" suit... 4-32
blind flying curtain assembly .. 4-32
checklists .. 4-32
emergency signal system. 4-32
map and data cases .. 4-32
miscellaneous pares storage .. 4-32
rear view mirrors .. 4*32
relief tubes .. 4-32
windshield wiper ..4^31
Missed Approach
flight computer with IL$.. 9-13
go*around procedure ................... 9-13
Missile Launch Accumulator Air Gage. 5-7
Movement, Aileron and Rudder ................ .... 5*10
N
Navigational Equipment, Radio and .. 9*7
Navigation Set, VHF. 4-14
Night Flying .. 9-16
landing .. 9-16
takeoff . 9-16
Normal
landing .. 2-18
takeoff . 2*12
Normal Fuel Sequencing .. 7*4
Nose Wheel Steering Hydraulic System ..1*47
Nose Wheel Steering System .. 1*47
nose wheel steering button .. 1*47
O
Obstacle Clearance Takeoff. 2-13
Oil Pressure Gages. 1-9
Oil Quantity Data . 1-17
Oil Supply System. 1-17
failure . 3-21
pressure gages. 1*9
Omnirange and Radio Range Approaches .. 9-9
Operating Flight Strength Diagram ........... .5-11—*5-13
Operation, System and Equipment
afterburner .. 7-3
anti-icing .. 4*6
automatic approach equipment. 4-29
autopilot . 4-28
brake ........................................ 7-3
cabin air-conditioning. 4-2
canopy defogging .............................. 4-4
command radio.4-13
eyelid . 7-2
flight computer. 4-18
fuel system ... . 7-3
IFF . 4*19
interphone . 4-13
oxygen .4-25
radio compass .4-14
single-point fueling .. 4-31
speed brake 7-6
VHF navigation. 4 -16
wing flap ... . 7-6
Optical Sighthead. 4-29
Overheat Warning Lights and Test Switch, Fire and .... 1-52
Overspeeding at Altitude, Engine. 7-2
Overtemperature Versus Engine Life. 7-2
Oxygen Duration Hours Chart ..4-22
Oxygen Mask Connection.. 4-25
Oxygen Regulator Panel .. 4-23
Oxygen System . 4*22
emergency operation.4-25
normal operation 4-25
preflight check . 4-24
pressure gage and flow indicator.4-24
regulator .4-23
diluter lever .. 4*23
emergency lever... 4-23
supply lever .. 4-23
warning system switch and indicator lights.4*24
P
Panel Vibrators, Instrument. 1-48
Parking Brake Lever...... , ,, 1-48
Pattern, Approach to. 9-19
Pedals
brake . 1-48
rudder . I -37
Pilot's
canopy
handgrips. 1.55
(Boldface type denotes illustration)
7
Index
Pilot's—Stalls
T.O. IF-89H-I
Page
jettison ' f T ,T handle .. . r . .........,, „, , J-56
switch ....... . 1-54
cockpit lighting . *. .... t ^, 4-21
C-4 cockpit lights 4-22
rheostats . . ,.. 4-21
duties 8-1
microphone switches ...,,. .. 4-8
Pilot’s Center Pedestal. 1-13
Pilot's Instrument Panel .. 1-10
Pilot's Left Console . . ... , ..... . . 1-11
Pilot's Left Vertical Console ..... 1-12
Pilot's Miscellaneous Control Panel . . ..1-22
Pilot's Right Console ...... 1-15
Pilot's Right Vertical Console ... 1-14
Pitot Heat Switch .. 4-5
Position Lights and Control Switches ................ 4-20
Power Supply System, Hydraulic ..,. 1-33
Power Supply Systems, Electrical .. 1-23
Power System, External .. 1-25
Preflight Check... 2-1
airplane . 2-10
before entering cockpit .... 2-4
before exterior inspection .. 2-1
engine ------ .... 2-11
entrance .. 2-4
exterior inspection... 2-1
interior check .. 2-4
oxygen - ; .......4-24
Preparation For Flight ............................ 2-1
Pressure Gages
canopy ejector ........ . 1*54
hydraulic system ....... *.... 1-35
oil... 1-9
oxygen .. 4-24
Pressure Regulator, Cabin , .... 4-1
Prohibited Maneuvers .. 5-10
acrobatics . 5-10
aileron and rudder movement.. ... 5-10
inverted flight .. 5-10
landing ....... .... *.. 5-10
rocket,'missile firing ...........__........... 5-10
spins . ........ 5-10
Pumps, Booster_ ............ .. .... 1-17
Pylon Limitations... 5-9
Pylon Tank, Damaged Tip or 3-23
Pylon Tank Jettison System . 1-18
jettison limitations .... ..... 5-10
jettison button .. 1*19
R
Radar Observer's
canopy
handgrips _ 1-55
switch ....... 1-55
cockpit lighting . 4-21
C-4 cockpit lights......... 4-22
rheostats .... .... 4-21
duties .. 8-1
after takeoff climb ..... 8-2
before entering cockpit .. 8-1
before landing ..... ....... 8-3
before leaving airplane .... 8-3
before takeoff....... 8-2
during flight ..... 8-3
exterior inspection .. 8-1
ground tests 8-2
interior check .......... . 8-2
rear cockpit ....... . 8-2
on entering cockpit ... 8-1
microphone buttons ...... 4-8
Radar Observer’s Cockpit—Front View , ... 4-10
Radar Observer’s Cockpit—Left Side ... 4-11
Radar Observer's Cockpit—Right Side .... 4-12
Radio and Navigation Equipment ... 9-7
Radio Approach . 9-9
Radio Approaches ,, .... 9-7
Page
Radio, Command ...... 4-13
Radio Compass .... 4-14
controls ........ 4-14
operation ...... ... 4-14
Radio Compass Control Panel ... 4-14
Radio Magnetic Indicator ......... . . 4-15
Radio Magnetic Indicator 4-15
Radio Penetrations .. ...» f , # . 9-B
Radio Penetrations ... 9-7
Radome Anti-Icing System ... 4-7
switches ... (t 4-7
Rear View Mirrors.. 4-32
Receiver, Glide-Slope 4-17
Receiving Set, Marker Beacon 4-17
Regulators
cabin pressure ..... . 4-1
oxygen . 4-23
Relief Tubes . ...... 4-32
Restarting Engine in Flight.... 3-6
Restrictions, Flight.,.. 2-1
Right Armrest, Ejection Seat ... 1*56
Right Engine ..... ......... 2-9
inoperative ... 3-8
Right Hydraulic System .. 1-35
Rocket/Missile Firing...... 5-1D
Rudder .. 6-4
electrical trim knob ... ...... . 1-38
movement, aileron and.... 5-10
pedals ....... 1-37
adjustment crank. 1-37
trim switch . ..*... 1-38
Running Off Runway on Landing .. 3-19
Runway Overrun Barrier Operation (some airplanes) , . 3-19
s
Safety Belt Automatic Release ... 1-62
Safety Belt Automatic Release.. 1-60
Seat Safety Pins ..... . 1-59
Seats, Ejection .. 1-56
Servicing Diagram .. 1 -64—1 -65
Shoulder Harness Inertia Reel Lock Lever ............ 1-66
Sideslip Stability Augm enter System ... 1-40
emergency operation ...... 3-28
power sw-itch .. 1*40
Sighthcad, Optical .. 4-29
Signal System, Emergency .. ...4-32
Simulated Forced Landing .. 3-9
Simulated Single-Engine Flameout . ... 3-9
Single-Engine
flight characteristics .. 3-1
go-arou tid ........ 3-9
instrument letdowns and approaches .............. 9-13
procedure , .... 3-1
takeoff .,, ........ 3-9
Single-Engine Landing Pattern ... . . .3-10—3-11
Single-Engine Service Ceiling. 3-2
Single-Phase Inverter
switch ....... 1 -29
warning light . 1-32
Single-Point Fueling Panel. 4-31
Single-Point Fueling System . ... . 4-30
Single-Point Fueling System.... 1-19, 4-29
controls 4-29
operation . 4-31
Smoke and Fumes Elimination . . ..... 3-13
Speed Brake Lever ..... 1-42
Speed Brakes and Wing Flaps Hydraulic System.1-39
Speed Brake System.... 1-41
emergency operation .. 3-28
lever ........ 1-42
operation ......... 7-6
speed brakes . 6-4
Speed Range ... *... *... 9-7
Spins--------5-10,6 2
Stalls .. 6-1
accelerated . 6-2
compressor . 7-1
8
[Boldface type denotes illustration)
Changed 13 February 1959
T.O. 1F-89H-T
Index
Stall—Wing
Page
Stall Speed Chart ........ ... 6-3
Starting 5-1
afterburners at high altitude. ................ 7-3
and ground check ..... 4-18
engine .. 2-7
before starting . ....... 2 - 7 , 9-17
cold weather procedure ....... . ..... 9-17
left ..,____. .____ 2-8
right ..... 2-9
Starting and Ignition System.1-6
altitude start and starter test switches. 1-8
starter and ignition switches . 1-8
starting power switch ..... ..... 1-9
Steering System, Nose Wheel ...... 1*47
Stick
control .. 1-37
forces ......... 6-6
Stick Forces Chart # . 6 - 6 — 6-7
Stopping Engines ... 2-22
Storage, Miscellaneous Parts .................. ..... 4-32
Storm
approaching the ............................... 9-16
Straight and Level Flight .. 4-28
Surface Icing ... .... ..... 9-14
T
Tachometers .. 1-9
Takeoff . 2-12
aborted.... 3.3
after, climb .. 2-13
hot weather procedure .. 9-20
before .. 2-10
cold weather procedure .... ..... 9-18
desert procedure ... ..,., 9-21
radar observer's duties .. 8-2
cold weather procedures ........ 9-19
continued. 3-4
crossw'ind .. ..... 2-13
desert procedures ,... 9-21
hot weather procedures . . 9-20
instrument .. 9-4
minimum run .. 2-13
nisht. 9 .I 6
normal .. 2-12
obstacle clearance ..... * *,. * 2-13
single'engine ____ 3.9
Takeoff Procedure .. .3-13—3-13
Tanks. Damaged .. t . 3-23
main ..... ♦*♦*♦*..... 3-23
tip or pylon . 3-24
....*... 3-24
Taxiing, Before .. 2-9
cold weather procedures .. 9-18
Temperature, Ambient, Exhaust Gas Temperature Versus 5-7
Temperature Variation, Exhaust Gas ...... 7-1
Tests, Ground ..... 2-9 4-28
cold weather procedure ... . ... 9-18
radar observer's duties .. g _2
The Airplane .. . .. .. t , . . . j_j
armament . M
dimensions ..,,., 1-1
gross weight .. 1-1
Thermal and Electrical Anti-Icing Systems. 4-5
anti-icing switch.. ................. 4-5
anti-icing warning light..... 4 _£
Page
Operation .. 4-6
descent . 4-6
in flight ....... 4-6
landing .. .... 4-6
takeoff . .. 4-6
pitot heat switch. 4-5
wing anti-icing override switch .................. 4-5
Three-Phase Inverter
switch . 4 ........... .... 1-32
warning light .. 1-32
Throttle-Actuated Fuel Shutoff Switches ............ 1*19
Throttles ... . .... 1-2
friction lever .. 1-6
Throttles .. 1*8
Tip or Pylon Tank, Damaged . 3-24
Tip Tank Fuel Dump System.. ..... 1-18
dump button ..* * - - • 1-19
Tire Limitation .... ..... 5-9
Tire Pressure Chart ....... 5-9
Trigger, Catapult Firing ............. .. . 1-60
Trim . 6-5
Trimming Procedure, Autopilot .. ♦ * ♦. ♦ 4-28
Trim System, Flight Control .... *.. 1-37
Turbulence and Thunderstorms . ... ..... 9-15
approaching the storm ..... 9-16
Turns With Flight Computer ... 9-5
Typical Dive Recovery ... 6-8
¥
VHF Navigation Set . 4-14
controls ..... .. ...4-15
ground check . 4-16
indicators ......... 4 . 4-15
course ....... 4-15
radio magnetic ... 4-15
operation ............* . 4-16
for communications... 4-17
with localizer .. 4-17
with VAR _ 4-16
with VGR ... 4 -16
VHF Navigation Control Panel ... 4-T4
Vibrators, Instrument Panel .......... . 1-48
Voltage Check ...... 2-9
w
Warning Lights Dimming Switch ........... , 4-2 1
Weight and Balance .. «... 2-1
Weight Limitations ....„ 5*15
Wet or icy Runway Landing ... , ., 2-19
Windshield Heat System. 4-7
knob ........ *..,, 4-8
Windshield Wiper ....... 4-31
Wing Anti-Icing Override Switch ... 4-5
Wing Drop, High Airspeed... * 6-6
Wing Flop Lever ... * ... # ... „ 1-41
Wing Flap Operation..., ,., 7-6
limitations .......... 5.9
Wing Flap System ... ..... 1-41
emergency operation ...... 3-28
lever and position indicator.. 1-41
Wing, Fuselage, or Electrical Fire .. 3 -13
Wing Tanks
booster pump failure . ... ..... 3-22
damaged . t 3*24
(Boldface type denotes illustration}
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